What Is an Example of an Operational Load Requirement
Aerodynamic loads (or air loads for short) refer to the forces and moments caused by the asymmetry of pressure on the surface of the aircraft. Air loads include forces (e.g., wing lift and drag) and moments (e.g., wing twisting and bending). Their size depends on the weight of the aircraft, the load factor, its geometry and, in turn, the dynamic pressure. The overall size is defined based on aviation authority requirements, such as 14 CFR Part 23 and 25. However, local values depend on geometry. Consider two aircraft, A and B, with the same mass and wing area, differing only in aspect ratio (AR) and conical ratio (TR). Suppose aircraft A has the highest AR and the lower TR. For the reasons described in Chapter 9, The Anatomy of the Wing, it will produce a higher bending moment than Aircraft B. Aircraft are designed to meet different types of loads, aerodynamic, inertial and operational, as described below: For each new load condition S, such as a normal lifetime production operation, each strand undergoes a change in axial load ΔPi,S depending on operating conditions such as temperature or pressure change. The relevant mechanical properties for PTC are the lateral stiffness of the receiving supports and the torsional stiffness of the position of the manifold structure.
To measure these quantities, the structure is carefully subjected to an increase in force and/or torque well below the expected operating loads. The resulting strains and twists can be measured using high-precision inclinometers and linear/angular measurements. For each strand, calculate the final redistributed axial multichain load based on the current load step: the effects of operating loads on geometry can be evaluated directly by geometric measurements such as photogrammetry (PG). Alternatively, mechanical properties can be measured and the effects of operating loads on geometry can be calculated by static analysis. In the design phase, finite element simulations (FEM) are highly recommended to derive deformations from CAD models and operating load assumptions. The depth of water in the GoM is related to the distance from the shore and the type of structure needed to withstand environmental and operational constraints. As the depth of the water increases, the distance to shore increases approximately linearly (with the exception of some locations at the mouth of the Mississippi River, where the continental shelf is short), and as the land distance increases, so does the cost of transporting land bases. The need for a robust platform and higher insurance and communication costs increases with water depth, contributing to higher operating costs. Differences in water depth should result in differences in economic boundaries for all the same things. In the case of a sandwich panel, the low shear stiffness of the core leads to further shear deformation. The approach to include this effect can be found in Plantema (1966) and Allen (1969). For bolt support at both ends, the buckling load of sandwich panels with variation in material properties induced by the thermal gradient was calculated in Gu and Asaro (2005) as follows: Identify the static wellhead loads j = 1 to m to be applied in the current load step S: W1,S, W2,S, …, Wm,S.
In the case of composite panels, experiments have shown that folds occur when heated under operational loads considered safe according to design criteria without assessing fire damage (e.g., Gibson et al., 2004; Dao and Asaro, 1999). The latter reference suggested that global buckling was a fairly common failure mode for structural size panels. In Gu and Asaro (2005), buckling loading was derived from the theory of non-homogeneous elasticity to account for the variation of the material along the thickness by temperature gradients. For sandwich panels, Allen`s formulation for additional core deformation by shear (Allen, 1969) was included. The buckling solution is discussed here. Note that the term limit load refers to a limit below which the aircraft can only deform elastically. This means that after removing the load, the aircraft will bounce back to its original shape. Final load refers to a limit beyond which the aircraft can fail. Between the limit and the final load, the aircraft can undergo plastic deformation (permanent change of shape) while being safe to fly. The final load is 1.5× of the limit load. For the thermo-mechanical problem, the staggered approach can be used.
The temperature field is first determined from the thermal model in section 4.2. The temperatures obtained are then substituted in the degradation law of section 4.3 to find the degraded properties of the material that determine the buckling load from [4.19] and [4.22] with geometric properties. It should be noted that, since in such an approach, mechanical solutions are independent of thermal solutions, these buckling solutions, as well as the skinfold solutions discussed in the next section, are applied to the temperature range with or without chemical decomposition. In Gu and Asaro (2005) and Gu et al. (2009), predicted buckling loads were compared to data measured by Asaro et al. (1999) and Dao and Asaro using [4.19] and [4.22]. (1999). Testing was conducted with the UCSD multi-axis firing tester (Asaro and Dao, 1997; Asaro et al., 2009b; Dao and Asaro, 1999) to ensure that the panels display a one-sided fire while maintaining the flexibility of mechanical loads inside and outside the aircraft.
The properties of the materials and the dimensional lengths of the samples were given in the references above. Figure 4.6 shows the buckling loads of a single plate at certain heating times resulting from the actual temperature profiles and the estimated buckling load curve obtained by connecting the discrete points to straight lines.