Aircraft Conceptual Design Lecture 6 PDF

Aircraft Conceptual Design Lecture 6: Structures and Loads Definitions Limit load (or applied load): The largest load the aircraft is expected to encounter. In the figure, limit load on the wings occurs during an 8g maneuver. Design load (or ultimate load): The highest load the aircraft is designe...

Aircraft Conceptual Design Lecture 6: Structures and Loads Definitions Limit load (or applied load): The largest load the aircraft is expected to encounter. In the figure, limit load on the wings occurs during an 8g maneuver. Design load (or ultimate load): The highest load the aircraft is designed to withstand. For safety reasons, aircraft are designed for loads higher than limit loads. Safety factor: The multiplier used on the limit load to determine the design load. Usually assumed as 1.5. For the plane in the figure design load should be 12g. 2 Definitions (load factor) Load factor (n): expresses maneuvering of an aircraft as a multiple of gravitational acceleration (g=9.81 m/s2) . Table below shows typical load factors. At level flight, load factor on an aircraft is n=1 (see figure on next slide). During flight, there will be extra loads on the aircraft, up to 9 times the level flight conditions for fighters. 3 Definitions (Load factor) L n 1 W 4 Definitions (V-n diagram) V-n diagram: depicts aircraft limit load factor as a function of airspeed. A V-n diagram is shown below. This diagram is used in Federal Aviation Regulations (FAR) part 23 or 25 for aircraft structural regulations At level flight, max. lift load factor is 1.0 at stall speed, meaning aircraft can not fly below that speed. It can be stalled at higher speeds through maneuvers such as steep turns. 5 Definitions (V-n diagram) At low speeds the maximum load factor is constrained by aircraft maximum CL. At higher speeds the maneuver load factor may be restricted as specified by FAR Part 23 or 25. • The maximum maneuver load factor is usually +2.5 . If the airplane weighs less than 50,000 lbs., however, the load factor must be given by the FAA (Federal Aviation Administration): n= 2.1 + 24,000 / (W+10,000) n need not be greater than 3.8. This is the required maneuver load factor at all speeds up to Vc, unless the maximum achievable load factor is limited by stall. • The negative value of n is -1.0 at speeds up to Vc decreasing linearly to 0 at VD 6 Definitions “High AOA” marks the slowest speed at which the max. load factor can be reached without stalling. At high AOA, direction of L is forward of body vertical axis (perpendicular to the wing in the figure) The forward component of the L creates an extra load on the wing structure. Dive speed (Vdive) represent the speed at max. allowed dynamic pressure. Structural limit of the aircraft is determined by the point “Max q” . For subsonic aircraft, Vdive is about 50% higher than level flight cruise speed. For supersonic aircraft, it is M0.2 higher than level speed at engine max. continuous power 7 Definitions Ve stands for equivalent airspeed; the speed at sea level that would produce the same incompressible dynamic pressure as the true airspeed at the altitude at which the vehicle is flying. That is, 2 2 q  1  SLVequivalent  1  altitudeVactual 2 2  Vequivalent   altitude  SL Vactual Ve stays constant w.r.t. dynamic pressure at any altitude. Gust loads are the loads acting on the aircraft under a strong gust or turbulence. Figure below depicts gust velocity U coming from under the aircraft, having a initial velocity V, at an AOA α. The change in AOA, Δα, would be small, so it can be approximated as: 8 Gust loads The change in total lift and load factor would then be Gust does not affect the aircraft instantly. The slow action of the gust is accounted for by using a statistical gust alleviation factor, K. Design requirements for gust velocities are derived from equivalent airspeed, hence actual effect of the gust is calculated using equivalent airspeed of the gust, Ude : 9 Gust loads Mass ratio stresses that small, light airplanes are affected quicker than larger ones. Also, Δn equation shows that load factor difference due to gust is more if the aircraft has low wing loading. Expected gusts are reduced at high altitude (more than 20,000ft). Vertical gust speed Ude is specified as shown below-left. A V-n diagram showing the gust load factors is given below-right. Vg is the max. turbulence speed. Drawing the V-n diagram for gust loads So, to construct the V-n diagram at a particular aircraft weight and altitude, we start with the maximum achievable load factor curve from the maneuver diagram. We then vary the airspeed and compute the gust load factor associated with the Vg gust intensity. The intersection of these two lines defines the velocity Vg. However, if Vstall ng  Vg then we can set Vg  Vstall ng and use the maximum achievable load at this lower airspeed. Next we compute the gust load factor at VC and VD from the FAA formula, using the appropriate gust velocities. A straight line is then drawn from the VB point to the points at VC and VD. 11 Combined V-n diagram In this figure, the V-n diagrams from maneuver and gust considerations are combined. Since the gust loads are greater than the assumed limit load, the assumed limit load may be raised to the dashed lines as well. With the safety factor of 1.5, the design load factor would then be raised accordingly. 12 Using V-n diagram data Information from the V-n diagram is used to calculate loads on the lifting surfaces. Critical loads occur at velocities of High AOA and Max q, and where the gust load factor exceeds assumed limit load factor. Once the Lifting force and pitching moment formed on the wing is determined, required lift on the horizontal tail to balance the wing pitching moment at critical conditions is calculated. Using the total lift on the wings and tail, spanwise and chordwise load distributions are determined. Wind tunnel or aerodynamic analysis codes may be utilized. Once the spanwise load distribution is known, wing or tail bending and torsion stresses can be calculated. 13 Material selection Factors considered for finding the best material include Yield and ultimate strength Stiffness Density Fracture toughness Fatigue crack resistance Creep Corrosion resistance Temperature limits Producibility Repairability Cost Availability 14 : (see figure) : resistance to deformation under force Material selection Fracture toughness: the total energy per unit volume required to deflect the material to the point of fracture. It is equivalent to the area under the stress-strain curve. Creep: tendency of materials to slowly and permanently deform under low but sustained stress. Plastics, composites and some Ti alloys are subject to creep at room temperatures. Corrosion occurs when materials are exposed to moisture, salt, aircraft fuel, oils, battery acid, engine exhaust, etc. • Corrosion products at the component surface tend to form a protective coating that delays further corrosion. However, if the component is subjected to a tension stress, the coating cracks and corrosion resumes inside the component. Therefore corrosion is accelerated when materials experience sustained stresses. “Stress corrosion” can cause fracture at a stress level 1/10th the ultimate stress. Residual tension stresses in materials are not desired. Producibility indicates ease of fabrication. Generally, if the material is strong, stiff etc., it is more difficult to form desired shape and repair. Material availability increases with the possibility of obtaining the material when desired. 15 Common aerospace materials Wood: Used in all early aircraft. Not so popular any more. Still used in homebuilt aircraft. • Cheap, easy to shape but requires craftsmanship • low strength and stiffness but good strength-to-weight ratio • low resistance to moisture, insects Steel: Used in airframes, especially in early aircraft. • Very high strength and stiffness (depending on alloy), but heavy • Cheap, easy to fabricate • High resistance to fatigue, temperature, corrosion (stainless steel) Aluminum alloys: The most common material in modern aircraft. • Relatively low cost, • high strength and stiffness, light weight • low fatigue properties, • Good corrosion resistance • Low resistance to high temperatures 16 Common aerospace materials Titanium: Very good aerospace material. • Strength to weight ratio and stiffness better than Al • Resistant to heat, corrosion. • Very expensive , hard to fabricate, • low availability Mg alloys: Used in engine mounts, wheels, wings, control hinges etc. • Good strength to weight ratio • low cost, easy to form • Low resistance to corrosion (must have protective finish) • Flammable High-temperature Ni alloys: suitable for hypersonic and re-entry vehicles • Can operate at extremely high temperatures • Heavy • Difficult to form 17 Common aerospace materials Other non-metallic materials: include reinforced plastics, rubber , sealants, cockpit transparencies. • Older reinforced plastics have poor stiffness, sensitive to temperature, hard to attach to metals. • Newer types include composite materials. They have good strength, light weight, but are expensive to manufacture. They are getting more popular in modern aircraft. Composites are mostly a reinforcing material (e.g. fiber) suspended in a matrix material that stabilizes the reinforcing material and bonds it to the adjacent reinforcing materials. Whisker-reinforced composites have short strands of fibers randomly distributed in the matrix. Filament-reinforced composites have higher strength and tailored to withstand loads in expected directions. 18 Composites Fiber orientation can be tailored as shown. Tailoring leads to great weight savings, but the resulting material is anisotropic; i.e. the material can only carry the loads at a number of directions. Metals and whisker-reinforced composites are isotropic; they can carry the same load at any direction. By having several layers of orientation, it is possible to carry loads form many directions. 19 Composites A sandwich construction using composites is common in aircraft. It is a structural sandwich of two face sheets (typically Al, fiberglass-epoxy, or graphite epoxy) and a honeycomb or solid foam core in between. Face sheets carry tension and compression loads due to bending. Core carries shear loads and loads perpendicular to the faces. 20 Composites Although composites have great strength to weight ratio, they have several issues: • They can not carry concentrated loads; causes problems attaching components. • Material properties are sensitive to manufacturing process, temperature and ultraviolet exposure, and exact ratio of fiber to composite. Every part will have slightly different properties. • Mild damage in composites occur internally, and it is impossible to inspect. Therefore they must be designed to carry limit loads while damaged. Properties of some composites materials are given in Table 14.4 in your textbook. 21