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Introduction of Advanced Aerospace Structures Assignment
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The advanced aerospace structure has covered the area of research along with the aerodynamics, combustion, dynamics, flight mechanics power of the aerospace and propulsion. It is one of the primary research of the department which includes mechanical engineering and also aerospace engineering. In this study, it has been described the efforts of this research and calculations about aerodynamics and its flow control, combustion, the reacting flows of turbulence, mechanics of flight, dynamic control of this aerospace structure, structures, and material, and many more calculations about this.
Question 1
- The bending moment and shear force calculation of the wing of an aerospace structure and Schrenk Distribution,
The relation of loading and deflection,
The wing of the aerospace structure, the load of the beam is distributed along with the span,
q(y) = L’(y) - N g m’(y)............................(1)[ the q stands for loading of the net beam, y stands for coordinates with spanwise, L’ stands for distribution or Schrenk distribution, N stands for the factor of load, g stands for the gravitational acceleration and m’ stands for the mass of the wing along with the span distribution]
From the above equation, the resultant of the distribution has produced a angle of deflection which is denoted as θ(y) and there is also the deflection w(y) of this beam of the wing,
With the help of the standard equation of differentiation, it has derived that the primary structure of this model is along with the axis of y-direction, and it is also related to the loads and the deflections for the loads q(y) and the stiffness of the bending EI(y).
dS = q dy………………………………(2)
dM = S dy………………………………..(3)
dθ = M/EI . dy……………………………(4)
dw = θ dy…………………………………(5)
It is too much necessary to integrate the boundaries and their conditions. For the cantilevered beam of the wing,
y = b/2 : S = 0……………….(6)
y = b/2: M = 0……………….(7)
y = 0: θ = 0…………………(8)
y = 0: w = 0……………….(9)
The lift and the lift coefficient is determined, with the help of the above boundaries and angles it has been calculated,
The force of the wing, D = 1/2 CD ρ v2 S
The angle of the deflection ∝ = CL / (dcL / da) + ∝CL=0
Then the calculation,
dcL / d∝ = 0.1 AR / AR + 2 = 0.1 . 7.7 / 7.7 + 2 = 0.079 / degree
Shear Force and Bending Moment of the Aerospace Structure
(Source: Otsuka and Makihara, 2019)
The Schrenk spanwise distribution,
The Schrenk method has relied on the matter of fact that the distribution of the lift across the wing of the aerospace distribution which doest does not vary much from the elliptic, whether itb is highly a span wise distribution. This Schrenk process has required so much elliptic spanwise distribution platform along with the semi span of the wing of aerospace structure and after that it will be modified by considering the chord of the wing and its variations.
The method of the Schrenk distribution’s calculation,
The lift of the load elliptical,
L’elliptical = 4S / πb √ 1 - (2y / b)2……………………………(1)
The lift of the load Planform,
L’planform = 2S / (1 + λ)b. (1 + 2y / b (λ - 1))..................................(2)
And it has been calculated as Schrenk lift,
L’Schrenk = L’elliptical + L’planform / 2………………………………..(3)
Span Wise Distribution of aerospace structure
(Source: Otsuka and Makihara, 2019)
- The calculation of the lift coefficient of the Aerospace structure and How to prevent from stalling of the Aerospace structure,
Calculation of the Lift coefficient of the aerospace structure,
In the beginning, loads of the aerodynamic structure on the aircraft are to bring the data of the airfoil. The airfoiled has been scaled as the unit chord as mentioned in the table and also plotted. The wing of the aerospace wing has 15 percent of thickness and also 6 percent camber.
There are some estimated data for the lift curve and the slope for this airfoil of
Α = 0.095 / degree and the zero lift is towards the negative value of 8 degrees. For calculating the lift load on the upper and lower portion of the wing of the aerospace structure,
- C upper = Cla (AR / AR + 2) (α - α0L)
= (0.095) ( 5.14 / 5.14 + 2) [ 0 - (-8)]
= 0.55
CL.lower = Clα (AR / AR + 2) (α - α0L)
= (0.095) ( 5.67 / 5.67 + 2) [ 1 - (-8)]
= 0.63
The equation of the lifted load which has been utilized for calculating the force applied to the wings of the aerospace structure,
Lu = q SCL
= (0.277 psi) (20418.3 in2) (0.55)
= 3110.73 lb
LL = q SCL
= (0.277 psi) (12630.72 in2) (0.63)
= 2204.19 lb
or , the balance of the moment is determined to calculate the force and coefficient is necessary for the aircraft balance,
Ft = LuXu - LLXL / XT
= [(3110.73) (11.60) - (2239.17) (10.41)] / 190.12 = 79.40 lb
Lift Coefficient of the wing
(Source: Adam et al. 2018)
There is a method for prevention of stalling of aerospace structure, as the determination of the target along with the speed on the air is to slow down the aircraft for the warning of stalling which has been desired to slow the configuration of the flight, pitch the nose towards the downward in the slightest for the estimation in order to warn the stalling and add the power for keeping the altitude and the note for the airspeed. There is another way to reduce the stalling for the aircraft such as the contamination of the wing of the aerospace like frost or the ice can be able to minimize the amount of the lifting load which has been produced by the wing and also raise the speed of the stall. The several changes along with the geometry from the high lifting devices like the leading edge of the aircraft are to increase the coefficient of the lower speed of stalling (Maymon, 2021).
- Calculation of the ultimate shear force of the wing and discussion about the location where it is minimum and maximum
The ultimate shear force theory has stated that the breaking down of the material of the aircraft has depended on the maximum shear force which is attained in the element. It has been assumed that the starts of the yielding in the planes of the shear force which is been maximized. According to this theory, the maximum shear stress at a point has been reached along with the strength of the wing.
Here is the calculation,
d = 2m, t = 20 mm, p = 1.5 N/mm2 and F = 2500 kN
The solution of this,
Ultimate shear force,
σn = σxcos2θ + σy sin2θ + txy sin 2θ………………(1)
σx = pd / 4t = 1.5 x 2 x 103 / 4 x 20 = 37.5 N / mm2
σy = pd / 2t = 1.5 x 2 x 103 / 2 x 20 = 75 N / mm2
σx = F / πdt = 2500 x 103 / π x 2 x 103 x 20 = 19.9 N / mm2
The Ultimate Shear Force
(Source: Carrera et al. 2018)
- Calculation of the ultimate bending moment of the wing of the aerospace and discussion about the location where it is maximum or minimum
The maximum bending moment has been considered as the supported beam with the bearing of a lifted load, the maximum bending moment of the beam has occurred at the point on the maximum stress when it fails at the last moment. The airfoiled has been scaled as the unit chord as mentioned in the table and also plotted. The wing of the aerospace wing has 15 percent of thickness and also 6 percent camber (Kopsaftopoulos et al. 2018).
The force of the wing, D = 1/2 CD ρ v2 S
The angle of the deflection ∝ = CL / (dcL / da) + ∝CL=0
Then the calculation,
dcL / d∝ = 0.1 AR / AR + 2 = 0.1 . 7.7 / 7.7 + 2 = 0.079 / degree
The ultimate bending moment of the aircraft,
Mmax ≈ L / 2 x yMGC = nult W / 2 x yMGC
In this equation, L is a lifted load, and the nult is the ultimate factor of the load, and the W stands for the weight of the airplane,
It has been outcome as
Mmax ≈ nult W / 2 x yMGC = nult W / 2 x √AR x S / 6 ( 1 + 2 λ / 1 + λ) = nult W √AR x S / 12 ( 1 + 2 λ / 1 + λ)
Maximum Bending Moment
(Source: Kecskemety et al. 2022)
Question 2
calculation of the maximum deflection,
For the optimization of the maximum deflection, here is the calculation of each of the coordinates along with the wing of the aerospace structure,
k (y) ≡ d2w / dy2 = dθ / dy = M (y) / EI (y)
Here at the constant ratio from the aerospace structure which has been located in the wing of the aerospace structure,
The root at the wing y = 0
k (y) ? k0 = M0 / EI0……………………………..1
θ(y) = ?y - 0 k0 dy = k0 y…………………………2
W (y) = ?y - 0 θ dy = 1 / 2 k0 y2…………………….3
For the straight wing ct / cr = λ, the distribution of the chord is,
c(y) = Swing / b . 2 / 1 + λ [ 1 + ( λ - 1) 2y / b]
or, q(y) ? kq c(y) = NWfuse / b . 2 / 1 + λ [ 1 + ( λ - 1)?]
? ≡ 2y / b
S (y) = ? b/2 - y q(y) dy
S (?) = b / 2 ?1 - n q (?) d?
= NWfuse / b . 2 / 1 + λ . b / 2 ? 1 - n [ 1 + (λ - 1)?]
= NWFuse / b . 2 / 1 + λ . b / 2 [ 1 - ? + (λ - 1) . 1 / 2 (1 - ?2)]
M (y) = ? b/2 - y S(y) dy
M (?) = b / 2 ?1 - n S (?) d?
= NWFuse / b . 2 / 1 + λ . b / 2 [ 1 - ? + (λ - 1) . 1 / 2 (1 - ?2)] d?
= NWfuse / b . 2 / 1 + λ . b / 2 ? 1 - n [ 1 + (λ - 1)?] + [ 1 - ? + (λ - 1) . 1 / 2 (1 - ?2)] d?
Maximum Deflection of the Wing
(Source: Deane et al. 2019)
Balsa Wood - balsa is the softest and the highest timber which is used in commercial. It has been processed an unusual in a higher degree of the buoyancy and it has been provided an insulation at its higher efficiency against the sound and the heat. The location where these type of properties are more essential for the woods which has been adapted to a great number of utilization at its specialty (Deane et al. 2019).
Question 3
The boundary of the aircraft structure which has been contributed to the production along with the stress from the inside of the system,
The taken moments along with the axis throughout the center of the element which is parallel to the axis of z
txy δyδz δx / 2 + ( txy + δtxy / δx . δx) δyδz. Δx / 2 - txy δxδz . δy / 2
- ( tyx + δtxy / δy . δy) δxδz . δy / 2 = 0
Or,
t xy δyδzδx + δtxy / δx . δyδz . (δx)2 / 2 - txy δxδzδy - δtxy / δx . δyδz . (δx)2 / 2 = 0
The fuselage bulkhead
(Source: Li et al. 2021)
Conclusion
The resources which are technical to its problem not be separated from the air force and the other duties. It has been affected by the decision of the air force about the current and the several missions for the future. It is too much necessary for commercial aircraft which is supersonic in very different terms. The configuration of the aircraft has been reduced the sonic boom of intensity and there is so much information about the shock waves from the wings and also manage to improve the performance which is aerodynamics can be achieved laminar flow with this type of configuration.
References
Adam, T.J., Liao, G., Petersen, J., Geier, S., Finke, B., Wierach, P., Kwade, A. and Wiedemann, M., 2018. Multifunctional composites for future energy storage in aerospace structures. Energies, 11(2), p.335.
Carrera, E., Cinefra, M., Filippi, M., Pagani, A., Petrolo, M., Zappino, E., Garcia, A., Kaleel, I., Manish, N., Li, G. and Guarnera, D., 2018. ADVANCED NUMERICAL METHODS FOR FAILURE ANALYSYS OF METALLIC AND COMPOSITE AEROSPACE STRUCTURES.
Deane, S., Avdelidis, N.P., Ibarra-Castanedo, C., Zhang, H., Nezhad, H.Y., Williamson, A.A., Mackley, T., Davis, M.J., Maldague, X. and Tsourdos, A., 2019. Application of NDT thermographic imaging of aerospace structures. Infrared Physics & Technology, 97, pp.456-466.
Kecskemety, K.M., Ita, M., Rumreich, L., Cartwright, E. and Staniak, C.M., 2022. Using Concept Maps for Assessment in an Aerospace Structures Course: A Comparison of Scoring Techniques. In AIAA SCITECH 2022 Forum (p. 0576).
Kopsaftopoulos, F., Nardari, R., Li, Y.H. and Chang, F.K., 2018. A stochastic global identification framework for aerospace structures operating under varying flight states. Mechanical Systems and Signal Processing, 98, pp.425-447.
Li, Y., Xiao, Y., Yu, L., Ji, K. and Li, D., 2021. A review on the tooling technologies for composites manufacturing of aerospace structures: materials, structures and processes. Composites Part A: Applied Science and Manufacturing, p.106762.
Maymon, G., 2021. New Approach to the Reliability Verification of Aerospace Structures. Modern Trends in Structural and Solid Mechanics 3: Non?deterministic Mechanics, pp.77-94.
Otsuka, K. and Makihara, K., 2019. Absolute nodal coordinate beam element for modeling flexible and deployable aerospace structures. AIAA journal, 57(3), pp.1343-1346.
