Page 22 range to be detected. AE stress wave sources are associated with breaks in molecular structure, i.e. in polymers between main-chain linkages or weak secondary linkages.
The waves have a high frequency content (100 kHz – 2 MHz) which makes this technique insensitive to mechanical vibrations usually generated by the engines and other aircraft parts. As a crack propagates AE is generated and so, particularly for composite materials, the growth of flaws like delamination or cracks in the matrix or fibres can be detected before they become dangerous [31].
8. Morphing structures
Advances in composite material research and further study of failure modes in composite structures will lead to a new breed of aircraft which can heal themselves and also perform multiple roles by altering the shape of their components thus changing their aerodynamic properties and mission specifications. Proof of concept was demonstrated by Duenas et al that a low volume-fraction (5-10%) of magnetic particles is sufficient for enabling self healing of an approximate 150 micron x 5000 micron crack in a mendomer polymer using inductive heating. It was also demonstrated that carbon-fibre-composites can be fabricated to morph using an apparent shape memory effect of the same mendomer that was used to demonstrate the self-healing [25].In their paper Duenas et al describe a self healing system which can automatically heal its cracks without the requirement of an external sensing system or actuator. According to them, the autonomous crack healing is accomplished by dispersing microspheres containing a healing chemical called dicyclopentadiene
Page 23 (DCPD) and a polymerizing agent known as Grubbs’ catalyst. When a crack is initiated in the material, the high stresses associated with it cause the nearest microspheres to break, releasing the chemical, which after interacting with the catalyst, initiates a chemical polymerization reaction of the DCPD that heals the crack. Similarly fibres storing healing resin have also demonstrated by Pang et al, where when fractured, the resin flows into the damage sites within the structure. However the research into these carbon compounds is at a very early stage and some drawbacks still exist such as the catalyst and the healing agent degrade at high temperatures, at low temperatures their response time becomes slow and once the microspheres burst they can’t be reused thus the crack can be healed only once at a particular location.
Figure 8:Healing Cracks
[27]
Many engineering ideas came from observing nature; aircraft themselves were envisioned by observing nature. When Animal tissue is damaged blood flows out which clots and is also sensed by the brain which sends signals to increase the body temperature. Precisely this can be accomplished if the research done by Zako & Taka is combined with, one of the structural health monitoring systems described above.
According to Zako & Taka a polymeric material which hosts a second solid-state
Page 24 polymer phase can migrate to the damage site under the action of heat thus healing the crack [26].Biomemetic self healing i.e. healing mimicking nature can be encapsulated by a table prepared by Trask et al.
[27]
9.Flutter in Composite Wings and need for Vibrational Analysis
Aeroelasticity is a phenomenon which causes great instability in aircraft, vibrations in the wing causes flutter. Emergence of flutter compromises not only the long-term durability of the wing structure, but also the operational safety, flight performance and energy efficiency of the aircraft, Flutter in a wing causes its tips to rise and fall which will change the angle of incidence , thus resulting in instability [35]. The aeroelastic analysis of laminated composite wings is also vital to the prevention of failures induced by oscillatory motion. The aeroelastic instabilities will change, however, when a crack has initiated in a wing structure and must be accounted for by adjustment to the structural and dynamic model [36].Therefore Flutter not only results in aerodynamics
Page 25 instability it also causes crack initiation and propagation in carbon fibre composite wings. Taking all this into account it becomes apparent that vibrational analysis of composite wings is necessary for a safe aircraft design.
Page 26
Methodology and Results
The aircraft design was made using Advanced Aircraft Analysis (AAA) software and the analysis of the wing structure was done using Abaqus version 6.9.
1.Aircraft Design
:Aircraft design has now become an iterative process; therefore no new aircraft is built from scratch. A base aircraft is taken and improvements are made on its design depending on the mission specifications. Therefore for the purpose of designing a
Page 27 completely carbon fibre composite aircraft the Learjet 85 was taken as the base aircraft which is the world’s first completely CFC aircraft due to enter production in late 2012.Data of other similar aircraft designs for the iterations to be carried out in AAA (Advanced Aircraft Analysis Software) was found from a number of sources. The more the number of similar aircraft, the more accurate the iterations would be especially in the weight sizing module of AAA. Therefore a number of similar designs were researched and the aircraft solution that came up in this MSc thesis was based on the Learjet 85 but is a new design since all the data was not be available and was calculated in AAA from data that was available.
An initial sketch of the Learjet design drawn in AutoCAD gave a rough idea of the design parameters such as wing span, fuselage length etc.
(1)
Page 28 Most of the parameters changed as the design process progressed, however this rough sketch was extremely useful to keep the final design as close as possible to the original idea.
Weight Estimation: The weight of the aircraft determines all other aspects of the design such as the wing span, because the lift that the wings are required to produce will depend directly on the weight it has to lift, this in turn will affect the geometry of the control surfaces and other components. Therefore it is very important to estimate the weight of the aircraft depending on its mission specification. The aircraft designed in this MSc thesis is a midsized business jet and its mission specifications are given below:
No of passengers: 8 Max Cruising Speed: 470 Knots
Crew: 2 Specific fuel consumption (sfc) = 0.5 lb/h/lb
No of Engines: 2 (Turbo Fan Jet Engines) Range: 2500 nautical miles
After the mission specifications were finalized the flight segments were defined and then created in AAA.
Page 29 (2)
The software contains in-built typical values of “Mff” i.e. fuel fractions required to calculate the weight of the aircraft in each segment. However for the cruise and climb segment the software requires manual input based on the mission specifications.
Page 30 (3)
In the cruise segment the range, cruising velocity and fuel consumption were based on the base aircraft i.e. the Learjet 85.The lift to drag ratio was estimated from coefficient of lift (Cl) value of the wing from the equation Lift (L)= ½ p V^2 S Cl and from the drag coefficient (Cd),then a typical value of l/d was chosen from the Roskam Tables in AAA.
This value however changed when the aerodynamics module was completed and the values had to be adjusted till the required range, cruising velocity and fuel consumption was achieved along with the design point.
In order to get the second curve for the design point a regression curve was plotted by finding similar aircraft in the same weight category as the required design.
Input Parameters
R 2500.0 nm V 470.00 kts cj 0.500 lb/hrlb L/D 12.31
Output Parameter
Mf f 0.8057
Fuel-Fraction in Cruise Segment: Flight Condition 1
Advanced Aircraft Analysis 3.12 Project 09/14/09 2:16 PM
Page 31 (4)
After the regression curve was defined the number of passengers and crew was entered into the program along with their estimated weights.
Page 32 (5)
The design point was finally achieved after some further adjustments in the aerodynamics module.
The program then gives the useful load as an output.
(6)
Page 33 (7)
Geometry:
Wing: The wing airfoil chosen for this aircraft design was the eppler 423 and the airfoil coordinates were obtained from the UIUC Airfoil coordinate database [29].This database contains coordinates for all known airfoils which can be converted to
‘afl’ format from ‘txt’ by simply renaming the file for use in AAA. The values for the wing geometry were balanced according to the results obtained in the aerodynamics module. Initial values were estimated then later adjusted after the performance module required changes in the aerodynamics which in turn affected the geometry.
Page 34 (8)
(9)
Fuselage: The fuselage is constructed by defining a series of concentric circles bounded by a square which determines the circularity. The more the number of circles the straighter the fuselage section will be. For ρ= A/B as shown in the figure below, A was calculated by using the Pythagoras theorem while B was chosen depending on how circular the section being created needed to be.
W ing Horiz ontal Tail Vertical Tail Canard V-Tail Angles
Ventral Fin Fus elage Landing Gear Airplane AeroPack Scale
W ing Ty pe Selection
Straight Tapered Cranked W ing Fuel Volume Flap/Aileron/Tab Chord Length
Select Input Parameters Combination
(X,Y,Z)f us Fuselage Coordinate System Nf
stations 16
Output Parameter
Coordinates Defined
Fuselage Table: double click for Cross-Section Dialog
Station x 1 2.6444 0.0000 0.6111 0.6111 0.0000 0.0000 -0.6111 0.6111 0.6111 0.6621 0.6111 -0.6111 0.6966 2 5.2888 0.0000 1.1000 1.1000 0.0000 0.0000 -1.1000 1.1000 1.1000 0.6552 1.1000 -1.1000 0.7103 3 8.5943 0.0000 1.6528 1.6528 0.0000 0.0000 -1.6528 1.6528 1.6528 0.6138 1.6528 -1.6528 0.7103 4 16.5275 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103 5 27.7662 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103 6 31.0717 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103 7 39.6660 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103 8 46.2770 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103 9 55.5324 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.6345 2.3139 -2.3139 0.6345 10 60.2300 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345 11 64.1267 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345 12 66.1267 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345 13 66.8123 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345 14 67.1235 0.0000 1.8000 1.8000 0.0000 0.0000 -1.8000 1.8000 1.8000 0.7103 1.8000 -1.8000 0.6345 15 67.5000 0.0000 1.2000 1.2000 0.0000 0.0000 -1.2000 1.2000 1.2000 0.7103 1.2000 -1.2000 0.6345 16 68.1000 0.0000 0.2000 0.2000 0.0000 0.0000 -0.2000 0.2000 0.2000 0.7103 0.2000 -0.2000 0.6345
Fuselage Geometry: Flight Condition 1
Advanced Aircraft Analysis 3.12 Project 09/14/09 4:12 PM
Page 36 (12)
Horizontal and Vertical Tail: The horizontal and vertical tail construction is the same as the wing construction. However care must be taken while choosing the vertical tail airfoil since unlike the wing, it is mandatory for the vertical tail to have a symmetrical airfoil otherwise there will be a lift force generated only on one side causing the
0.00 10.00 20.00 30.00 40.00 50.00 60.00 70.00 80.00 90.00 100.00
50.00
Page 37 (13)
(14)
Nacelles
The nacelles which cover the jet engines were designed using the nacelle coordinate system without defining the apex so that front end can be open. Defining the nacelles helped in calculating the CG in Class 1 weights.
(X,Z)apexn Apex is not included
(X,Y,Z)n Nacelle Coordinate System
Nnstations 7
Output Parameter
Coordinates Defined
Nacelle 1 Table: double click for Cross-Section Dialog Station x 1 0.5000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103 2 1.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103 3 1.5000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103 4 4.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103 5 5.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103 6 7.0000 0.0000 1.3220 1.3220 0.0000 0.0000 -1.3220 1.3200 1.3200 0.7103 1.3200 -1.3200 0.7103 7 11.2387 0.0000 1.3220 1.3220 0.0000 0.0000 -1.3220 1.3200 1.3200 0.7103 1.3200 -1.3200 0.7103
Nacelle Geometry: Flight Condition 1
Advanced Aircraft Analysis 3.12 Project 09/22/09 4:18 PM
Page 39
Loads
(15) The VN diagram obtained shows the manoeuvrable region of the aircraft in the green curve. Starting from the left of the green curve the top and bottom end points indicate the value of the load factor ‘n’ at the 2 stall speeds. To the right of the curve the top end point indicates the load factor at cruise and the bottom end point indicates the value of the load factor at dive speed.
An aircraft experiences aerodynamic loads induced by the pilot and loads induced by atmospheric turbulence. Pilot induced load limits are quantified in a manoeuvring V-n diagram. Gust loads that result from sudden wind gusts are calculated by forming a gust V-n diagram. An aircraft must be designed for both limit and ultimate loading. FAR
§25.301 defines a limit load to be the maximum load an aircraft is expected to see in service. Ultimate loads are limit loads multiplied by a factor of safety. The factor of
0.00 100.00 200.00 300.00 400.00 500.00 600.00 700.00 800.00 900.00
3.00
Page 40 following excerpts from FAR §25.305 explain the structural requirements for the two load categories:
§25.305 Strength and deformation.
(a) The structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.
(b) The structure must be able to support ultimate loads without failure for at least 3 seconds.
The Velocity to load factor plot was plotted with values calculated from other modules and the Veas was calculated from the formula Veas= ρ √V where V= true air speed and ρ is density of air at that altitude.
Aerodynamics
Lift: The lift for the wing and empennage group was calculated using typical values found in the Roskam tables which are in built in the AAA software. The values in the aerodynamics module are adjusted according to their effect on other modules. Since some of the values such as the range and estimated aircraft weight are constant, these can be used as a reference for adjusting the aerodynamic module.
Page 41 (16)
Drag:
The drag segment in AAA is similar to the lift segment, however since this aircraft is a carbon fibre composite aircraft typical values of skin friction could not be used and had to be researched.
W ing Lift Distribution: Flight Condition 1
Advanced Aircraft Analysis 3.12 Project 09/18/09 3:42 PM
Page 42 (17)
The above plot is an output after the drag segment is completed; it gives coefficient of lift vs. coefficient of drag for various aircraft conditions such as take off gear up or
0.0000 0.2500 0.5000 0.7500 1.0000 1.2500
4.0000
clean,M 0.0161 BDPclean 0.0621
M ission Profile Table M ission Profile Wbeginlb WFbegin
lb WFused
lb h ft Vkts CL CD L/D
Segment Input Input Input Input Input Output Output Output
1 W armup 22571.5 5905.2 225.7 0 0
2 Taxi 22345.7 5679.5 111.7 0 0
3 Take-off 22234.0 5567.7 111.2 0 300
4 Climb 22122.8 5456.6 345.3 43000.0 410 0.3052 0.0219 13.92
5 Cruise 21777.5 5111.3 4230.8 43000.0 470.00 0.2081 0.0188 11.05
6 Loiter 17546.7 880.5 290.0 40000 430 0.1906 0.0184 10.36
7 Descent 17256.7 590.4 172.6 15000 300
8 Land/Taxi 17084.2 417.9 136.7 0 0
L/D from W eights: Flight Condition 1
Advanced Aircraft Analysis 3.12 Project 09/18/09 4:05 PM
Page 43 The lift to drag ratio from weights is also given as an output in the Class 1 drag segment, these values affect the original values in the weight segment and might change the design point completely, and therefore they were adjusted accordingly.
Performance:
The objective in the performance sizing is to get a matching plot between various performance factors such as landing distance, maximum cruise speed and stall speed.
The values had to be adjusted in the various modules till all the curves passed through the same point as shown below.
(19)
Aeropack: After the design was completed in AAA the model was then exported to Aeropack software for a 3-D model of the design.
(W/S) TO lb
f t2 (T/W)
TO
0.00 25.00 50.00 75.00 100.00 125.00 150.00
1.00 Maximum Cruise Speed Landing Distance
TL = 0 deg F WTO = 22571.46 lb
Page 44 (20)
Page 45