User:Eas4200c.f08.nine.s/Lecture 1

Group nine - Week 1

Aircraft Design

Airplanes are composed of many different parts, the main structural ones being the wing and the fuselage. The wing and fuselage are made up of many basic structural elements, and those elements act like beams in a torsion member. One main difference between beams and aircraft structural elements is that aircraft loads are in the form of air pressure and concentrated loads such as in the landing gear and passenger seats. As an engineer, we have to decide if these local loads are not excessive enough to cause major deflections.

Aircraft Structure

The main structural purpose of the wing is to transfer the load to the fuselage. The airfoil shape of the wing is based on aerodynamics considerations and will not be discussed in this article, which focuses on the structural components of the wing. The wing acts like a beam and torsion member. The wing is composed mainly of spars and ribs. "The spar is a heavy beam running from spanwise to take transverse shear loads and spanwise bending. It is usually composed of a thin shear panel, or web, with a heavy cap or flange at the top and bottom to take bending." ^[1] There may be more than one spar in a wing, but in general one carries the majority of the forces on it, and is called the main spar. The wing ribs are flat structures that are used to take in-plane loads and reduce the effective buckling length of the stringers and as a result it increases their compressive load capability. Ribs are placed chord wise and are supported by the spars.

A combination of using spars with or without stringers, thin or thick wing skin, thin or regular airfoils and stringers being manufactured or not as an integral part of the skin depends on the design characteristics of the wing. It varies according to need and performance.

The skin of the fuselage bears the shear stresses due to the torques and traverse forces. It also carries the hoop stresses due to internal pressures of the aircraft.

The stabilator of an aircraft acts as a horizontal stabilizer and elevator. Stabilators are found on the back portion of the wings and allows the pilot to have a higher pitching moment.

Fuselage Designs

The fuselage is the center of an airplane, it is the main portion that holds pay loads and passengers. The fuselage has relatively small loads applied to it. Most of the forces acting on the fuselage are forces from the wing and landing gear reactions, and the pay loads.

Monocoque is a single shell aircraft and vehicle design. The word Monocoque is french where 'mono' means single and 'coque' means shell. This design uses the shell to support all of structural loading. This design technique is not used very often, because the shell has to be fairly thick to support the loading which makes the aircraft heavy. A preferred design method is the Semi-Monocoque.

Semi-Monocoque is similar to the Monocoque design having a single shell. However in addition it has stiffening braces. The purpose of the braces are help support the necessary loading. The advantage of this design is the weight. The shell does not have to be as thick therefore reducing the weight. Most Semi-monocoque designs have a thin shell supported by braces running axially called stringers and longerons. These stringers and then supported by transverse braces called frames or rings.

Landing Gear

The landing gear of an aircraft is what makes it possible for an aircraft to take off and land. The landing gear commonly consists of a high strength steel alloy because of the extreme stresses that are experienced during take off and landing. In general, wheels are the main component in the landing gear, but other gear such as skis or floats may be used.

Aircraft Materials

Aircraft materials must be of high strength and stiffness while remaining lightweight. The typical types of materials that are used tend to be steel alloys, titanium, aluminum, and fiber-reinforced composites. Aluminum is a very attractive metal for use as an aircraft material because it is very light weight and although not as strong as steel and titanium alloys, it is sufficient to withstand the shearing loads on the fuselage of an aircraft. The 2024-T3 aluminum alloy is a common material used for the fuselage and lower wing skins. These areas are prone to fatigue and cyclic tensile stresses and that is why a stronger alloy used. The upper wing, however, uses a 7075-T6 alloy which has a higher strength and lower fracture toughness. The upper wing experiences less fatigue and that is why it is used. Carbon fiber is a non-metallic material that is used in aircraft design due to its high strength and extremely low weight. "Fiber composites are stiff, strong, and light and are thus most suitable for aircraft structures." ^[2] Steel and titanium alloys see their use on the control surfaces because they are very strong metals.

Finite Element Method

The finite element method (FEM) is "a numerical technique for finding approximate solutions of partial differential equations (PDE) as well as of integral equations. The solution approach is based either on eliminating the differential equation completely (steady state problems), or rendering the PDE into an approximating system of ordinary differential equations, which are then solved using standard techniques such as Euler's method, Runge-Kutta, etc."^[3]

FEM is used often in most fields of Engineering. Partial differential equations are used in the designing of aerospace structures for example, and as a result the finite element method is used to simplify these functions for Aerospace Engineers. FEM is widely used for stiffness and strength visualizations, as well as reducing weight, material, and costs for certain projects. It can also be used to pin point where structures may bend or twist, and determine the distribution of stresses and displacements.^[4]

Stress and Strain

Stress and Strain are engineering terms that refer to a force per unit area, and the amount of deformation on an object caused by that stress respectively. Stress and Strain are proportional by a factor known as the "Young's Modulus". This quantity relates directly to an objects stiffness; the higher the modulus, the higher the stiffness.

The Equation $\sigma =E\epsilon$ is read as "Stress is equal to Young's Modulus times Strain. Stress and Young's Modulus both have the units of Force per Area. Strain is a dimensionless value that denotes how much a material has deformed by a ratio of deformed length over the original length.

In the figure to the right represents a stress-strain curve. The point on the graph denoted by the number 1, signifies $\sigma _{u}$ , the ultimate stress, this term represents the maximum stress possible for any ductile material. Point 2, represents $\sigma _{y}$ which represents the yield stress. This stress is the maximum stress that a material can withstand without being plastically deformed and can return to its original shape. Point 3 represents the point on the stress-strain curve where the material will fail or rupture. Region 4 represents the strain hardening region, and as discussed in the lecture is also known as the plastic hardening. Region 5 denoted the area where the material is softening or necking. To the left of point 2 known as the elastic region, the slope of that straight line is very important and is known as the Young's Modulus referred to earlier.

Example Problem

A rectangular box beam that has a thin-walled section is designed to endure bending moment 'M' and torque 'T'. The width and height of the box-beam are given by 'a' and 'b' respectively. The length of the beam is given by 'L = 2(a + b)'. Result to the figure below. Find the optimum ration 'b/a' that gives the most efficient section 'm = T' and

\sigma _{all}=2\tau _{all}

. The shear stress due to torsion is given by

$\tau =T/(2ab)$

Show Figure

Solution'
t << a
t << b
L = l(a + b)

Need to find the optimal cross section carrying the maximum bending moment 'M' and the maximum torque $\tau$ . It can be assumed that the shear stress distribution is uniform along the walls since the cross section walls are very thin. Naming the corners of the cross section A, B, C, and D, we obtain the following expression.

$T=T_{AB}+T_{BC}+T_{CD}+T_{DA}$

→ $T_{AB}=(b/2)*(va)=(1/2)\tau *abt$

where $v=\tau *t$

Solving similarly for $T_{BC},T_{CD},T_{DA}$ , the following equation is obtained

$T=2\tau *abt$

→ $\tau =(1/2abt)$

Note: the solution will be continued on next week's homework.

Contributing Team Members

The following students contributed to this report:
David Phillips:Eas4200C.f08.nine.d (talk) 03:20, 25 September 2008 (UTC)

References

^ C. T. Sun. Mechanics of Aircraft Structures 2nd Edition Pg.10 New York: John Wiley & Sons, 2006.
^ C. T. Sun. Mechanics of Aircraft Structures 2nd Edition Pg.16 New York: John Wiley & Sons, 2006.
^ "XJamRastafire" Wikipedia. 11 Sept 2008 <http://en.wikipedia.org/wiki/Finite_element_method/>.
^ "XJamRastafire" Wikipedia. 11 Sept 2008 <http://en.wikipedia.org/wiki/Finite_element_method/>.

[1] C. T. Sun. Mechanics of Aircraft Structures 2nd Edition Pg.10 New York: John Wiley & Sons, 2006.

[2] C. T. Sun. Mechanics of Aircraft Structures 2nd Edition Pg.16 New York: John Wiley & Sons, 2006.

[3] "XJamRastafire" Wikipedia. 11 Sept 2008 <http://en.wikipedia.org/wiki/Finite_element_method/>.

[4] "XJamRastafire" Wikipedia. 11 Sept 2008 <http://en.wikipedia.org/wiki/Finite_element_method/>.

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