Conceptual Design of a Carbon-fibre Composite Aircraft And Finite Element Analysis of the Wing
Anirudh Narayan MSc Aerospace Engineering Student id: 0800059 Project Supervisor: Dr. Giulio Alfano Course Director Advanced Mechanical Engineering Brunel University, West London Page 1
Contents Acknowledgements....................................................................................................................... 4 Abstract ......................................................................................................................................... 4 Introduction .................................................................................................................................. 4 Literature Review .......................................................................................................................... 6 1.Carbon Fibre Composites in Aircraft: ............................................................................................... 6 2.Materials used in Carbon Fibre Composite Aircraft ........................................................................ 7 3.Aircraft Structure: ...................................................................................................................... 9 4.Adhesive Bonding Of Aircraft Structures: ...................................................................................... 14 5.Delamination/Debonding Failure: ................................................................................................. 14 6. Finite Element Analysis: ................................................................................................................ 15 7. Structural Health Monitoring System ..................................................................................... 19 Comparative Vacuum Monitoring (CVM): .......................................................................... 20 Carbon Nanotube Network: ................................................................................................ 21 Acoustic Emission Sensor: ................................................................................................... 21 8. Morphing structures ............................................................................................................... 22 9.Flutter in Composite Wings and need for Vibrational Analysis ..................................................... 24 Methodology and Results ........................................................................................................... 26 1.Aircraft Design: .............................................................................................................................. 26 2. Structural Analysis:........................................................................................................................ 45 Conclusions: ................................................................................................................................ 66 Results ............................................................................................................................................... 66 Project Management ........................................................................................................................ 67 Further Work..................................................................................................................................... 67 Gantt chart .................................................................................................................................. 68 References: ................................................................................................................................. 69 Appendix : ................................................................................................................................... 73 Guidelines on interchanging between Abaqus and AAA .................................................................. 73 Email from Grob Aircraft ................................................................................................................... 74
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Table of Figures Figure 1:Manufacture of CFC fuselage.......................................................................................... 7 Figure 2:Fuselage Structure .......................................................................................................... 9 Figure 3:Wing Arrangement ....................................................................................................... 11 Figure 4:Wing Structure .............................................................................................................. 11 Figure 5: Equating SHM to the Nervous system ......................................................................... 19 Figure 6:SHM flow chart ............................................................................................................. 20 Figure 7: CVM sensor .................................................................................................................. 20 Figure 8:Healing Cracks ............................................................................................................... 23
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Acknowledgements I would like to thank my supervisor Dr. Giulio Alfano, Course Director Advanced Mechanical Engineering, who guided me from the initial title of the thesis to its final conceptualization and also taught me how to use Abaqus a software in which I had no previous experience . I would also like to thank Dr. Cristnel Mares, Course Director Aerospace Engineering who taught me about aircraft design and AAA software during my course at Brunel University.
Abstract In this MSc thesis under the guidance of Dr. Giulio Alfano a conceptual design of a carbon fibre composite aircraft was made and FEA was done on its wing. A procedure to design and analyze the structural components of an aircraft in Abaqus and optimize the design that was conceptualized in AAA was established. For this purpose the aircraft was designed in AAA software. The shell of the aircraft was then modelled in Aeropack and exported to Abaqus, where structural components were modelled and assembled into the wing. Vibrational analysis was then conducted to verify the structural integrity of the assembly and linear elastic analysis of the wing was conducted to verify the structural integrity at steady level flight during cruise.
Introduction The current economic conditions have resulted in the transfer of billions of dollars worth of wealth from the middle class to a few hundred elite. Thus it is necessary for the aerospace industry to adapt to the situation and find strategies to tap into this market of highly concentrated wealth. The practice of designing large transport jets is Page 4
already producing losses as air travel amongst the middle class is on a decline. The design of highly efficient business jets is one of the answers to tapping into this wealthy market which places more emphasis on aesthetics and comfort, unlike airline companies. Since mid-sized Carbon fibre Composite aircraft can be moulded into desired shapes and offers better strength to weight ratio, more aesthetic and aerodynamically efficient designs are now possible. However some components of the aircraft can’t be moulded as whole or bolted together as this can significantly reduce the strength and durability of the carbon fibre composite components. Therefore they have to be adhesively bonded together. A failure mode known as debonding occurs at these bond locations and delamination occurs inside the layers of carbon fibre. Since it is very difficult to detect delamination/debonding, it is important to incorporate safety features in the design phase itself. For example, the Eurofighter aircraft has a structural health monitoring system, if it is understood where delamination has the highest chance of occurring, piezoelectric material (material which develop a voltage difference on being deformed) can be added. This data can be integrated into the structural health monitoring system, if data about the rate of delamination/crack propagation is studied for a particular aircraft. A literature review on carbon fibre composite aircraft, delamination failure, Aircraft structures and Structural Health Monitoring systems was done to bring all these different fields of study together in order to have a better understanding and possible application to the carbon fibre composite aircraft being designed in this MSc thesis.
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Literature Review The literature review was conducted on a number of topics that were relevant to the dissertation. The reference material included published research material, books, lecture notes and websites.
1.Carbon Fibre Composites in Aircraft: Composites are materials which compose of two or more organic and inorganic materials. One material functions as a matrix and the other material functions as reinforcement. The most common matrix materials are "thermosetting" materials such as epoxy, bismaleimide, or polyimide. The reinforcing materials can be glass fibre, boron fibre, carbon fibre, or other more exotic mixtures [19].The main driver for using composites in aircraft is their high weight to strength ratio, which results in more fuel efficient aircraft. Another major advantage of using carbon fibre composites in aircraft is that they can be layered, with the fibres in each layer running in a different direction, therefore the designer can design components which behave in a particular way for example a component can be made to bend only in a particular direction. This behaviour resulted in the design of forward swept wing aircraft which would not have been possible with metals as they would bend during flight [19].Other advantages include part reduction, complex shape manufacture, reduced scrap, improved fatigue life, design optimisation, and generally improved corrosion resistance. The manufacture of CFC aircraft components is generally done in three steps, first a mould of the component is layered with composite material according to the design specifications. Once the component is laid-up on the mould is enclosed in a flexible bag tailored approximately to the desired shape and the assembly is enclosed usually in an Page 6
autoclave, a pressure vessel designed to contain a gas at pressures generally up to 1.5MPa and fitted with a means of raising the internal temperature to that required to cure the resin [20].Another method in which an expensive autoclave is not required is the vacuum method, in this method the space between the mould and the composite layup is evacuated of air and then it is heated to cure the resin. In the final step of manufacturing the mould is removed and if required, cuts are made in the component for example windows are cut into the fuselage.
Figure 1:Manufacture of CFC fuselage
[21] The main challenges restricting the use of CFC in aircraft are material and processing costs, damage tolerance, repair and inspection, dimensional tolerance and conservatism associated with uncertainties about relatively new and sometimes variable materials [20].
2.Materials used in Carbon Fibre Composite Aircraft Grob Aircraft, a company which builds carbon fibre composite aircraft and has won a contract from Bombardier for the design of the world’s first all composite business jet, the Learjet 85 was contacted for information about the various fabrics and resins they Page 7
use to manufacture their aircraft. The email from them is attached in the appendix. According to them, their aircraft are produced by use of wet lay-up composite materials, in this method the fabric or mat is saturated with liquid resin and the layup is obtained by building layer upon layer till the desired thickness is reached. Glass fibre is made by the following process, when quarry products (sand, kaolin, limestone, colemanite) are blended together at 1,600 degree Celsius, liquid glass is formed. The liquid is passed through micro-fine bushings and simultaneously cooled to produce glass fibre filaments from 5-24m in diameter. The filaments are drawn together into a strand (closely associated) or roving (loosely associated), and coated with a “size” to provide filament cohesion and protect the glass from abrasion [32]. The resins that are used in fibre-reinforced composites are sometimes referred to as ‘polymers’. All polymers exhibit an important common property; they are composed of long chain-like molecules consisting of many simple repeating units. Manmade polymers are generally called ‘synthetic resins’ or simply ‘resins’. Polymers can be classified under two types, ‘thermoplastic’ and ‘thermosetting’, according to the effect of heat on their properties, Thermoplastics, like metals, soften with heating and eventually melt, hardening again with cooling. This process of crossing the softening or melting point on the temperature scale can be repeated as often as desired without any appreciable effect on the material properties in either state. Typical thermoplastics include nylon, polypropylene and ABS, and these can be reinforced, although usually only with short, chopped fibres such as glass. Thermosetting materials, or ‘thermosets’, are formed from a chemical reaction in situ, where the resin and hardener or resin and catalyst are mixed and then undergo a non-reversible chemical reaction to form a hard, infusible Page 8
product [33]. Fibreglass is the most common composite material, and consists of glass fibres embedded in a resin matrix. To make a composite structure, the composite material, in tape or fabric form, is laid out and put in a mould under heat and pressure. The resin matrix material flows and when the heat is removed, it solidifies. It can be formed into various shapes. In some cases, the fibres are wound tightly to increase strength [34].
3.Aircraft Structure: Aircrafts are designed to play a variety of roles according to their mission specifications, however all aircraft generally have certain primary components i.e. Fuselage, Wing, Empennage, Power Plant and Landing Gear.
Fuselage:
Figure 2: Fuselage Structure
[3] The Fuselage’s primary function is to carry the pilot and the payload or passengers. Early fuselage designs had a box structure; the structural elements resembled those of a bridge, with emphasis on using linked triangular elements. The aerodynamic shape was completed by additional elements called formers and stringers and was then Page 9
covered with fabric and painted, however this kind of a structure proved to be heavy and modern aircraft use what is known as a semi-monocoque structure. In this type of an arrangement the skin is the main load carrying member. A series of frames in the shape of the fuselage cross sections are held in position on a rigid fixture, or jig. These frames are then joined with lightweight longitudinal elements called stringers. These are in turn covered with a skin of sheet aluminium, attached by riveting or by bonding [1]. Since carbon fibre composites can be layered over a mould they a full monocoque structure can be used. One important safety consideration to be taken in a carbon fibre composite fuselage is that unlike its all metal counterpart, a carbon fibre composite fuselage doesn’t provide shielding from lightning strikes. Some promising developmental lightning protection methods that should be considered are aluminium diverter strips, aluminium wire mesh, and aluminium flame spray [22].Some aircraft designs such as the ‘flying wing’ design used in the Northrop YB-49 Flying Wing and Northrop B-2 Spirit bomber do not have a separate fuselage ,the fuselage is a thickened portion of the wing. Conversely some designs use the fuselage as the lifting surface instead of the wing; Examples include NASA's experimental lifting body designs and the Vought XF5U-1 Flying Flapjack.
Wing: The wing of an aircraft provides the lift necessary for the aircraft to fly. The curved shape of the airfoil causes the air on the top surface to move faster than the lower surface which causes a difference in pressure resulting in lift. Three systems are used to determine how wings are attached to the aircraft fuselage depending on the strength of a wing's internal structure. The strongest wing structure is the full cantilever which is attached directly to the fuselage and does not have any type of Page 10
external, stress-bearing structures. The semi cantilever usually has one, or perhaps two, supporting wires or struts attached to each wing and the fuselage. The externally braced wing is typical of the biplane (two wings placed one above the other) with its struts and flying and landing wires [24].
Figure 3: Wing Arrangement
[24] To maintain the aerodynamic shape of the wing, it must be designed to maintain its shape even under extreme stress. The primary components of the wing which form its structure are the skin, stringers, ribs and spars.
Figure 4: Wing Structure [3]
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Skin: The primary function of the wing skin is to form an impermeable surface for supporting the aerodynamic pressure distribution from which the lifting capability of the wing is derived. Forces are transmitted to the ribs and stringers from the skin through plate membrane action [1]. Stringer: Although the thin skin is efficient for resisting shear and tensile loads, it buckles under comparatively low compressive loads, increasing the thickness is not an option because of the weight penalty. Therefore stringers are attached to skin and ribs thereby dividing the skin into small panels and increasing the buckling and failing stresses [1]. Spar: It is the main load carrying member of the wing. It resists shear and torsional loads also supports the skin, the spar flanges enabling them to support large compressive forces. Ribs: They maintain and support the airfoil shape of the skin. The other attachments on the wing which perform no structural function but are important from the aerodynamic point of view are flaps, ailerons and winglets which are used on some new wing designs. Flaps: They provide the extra amount of lift required at low speeds during manoeuvres like landing. Ailerons: They are used to roll the aircraft to one side during turning manoeuvres. When an aileron is deployed on one of the wings, more lift is generated on that wing thus rolling the aircraft in the opposite direction. Page 12
Winglets: These are found in new aircrafts as small vertical wings attached to the wing tips. Their function is to prevent induced drag i.e. vortices formed at the tips which cause drag.
Empennage: The empennage consists of the vertical and horizontal tails. The structure of the tails is similar to the wing. Vertical Tail: The vertical tail provides lateral stability and the attached rudder helps in the yaw movement of the aircraft. Horizontal Tail: The horizontal tail prevents the nose of the aircraft from pushing downwards due to the lift provided by wings by providing negative lift i.e. a force opposite to the direction of lift provided by the wings. In canard aircraft like the EuroFighter the horizontal tail is near the nose but gives lift in the same direction as the wings thus lifting the nose.
Propulsion: Aircraft primarily use propellers and jet engines for propulsion, some aircraft use ramjet engines which can function only at supersonic speeds for added propulsion. They are often used in conjunction with jet engines to achieve the right velocity to function. Propellers: The propeller blades are made in the shape of an airfoil, when the blades are rotated they produce lift which in this case results in thrust for the aircraft. Most propeller aircraft have propellers which pull the aircraft forward and are called tractor propellers. Aircraft which use the propeller to push to push it forward are known as pusher propellers. The engines used to rotate the propellers are piston engines. Page 13
Jet Engines: According to Rolls-Royce the jet engines work on the principle of ‘Suck Squeeze, Bang, and Blow’. Cold air from the atmosphere is sucked in by the fan which is then compressed by the compressor; fuel is then mixed with the compressed air and ignited this causes the hot high energy air to come out of the exhaust nozzle at high velocity which gives thrust to the aircraft.
Landing Gear: Most modern aircraft use retractable landing gear as there is a considerable increase in drag when they are deployed. Amphibious aircraft use floatation devices instead of wheels for landing.
4.Adhesive Bonding Of Aircraft Structures: Using adhesive bonding for joining composite parts provides many advantages such as low cost, high strength to weight ratio, low stress concentration, fewer processing requirements and superior fatigue resistance and environmental resistance[28].Since welding is not possible for carbon fibre composites and riveting makes them weak, adhesive bonding is the ideal method of joining CFC components. Adhesive bonding is used mainly for attaching stringers to fuselage and wing skins to stiffen the structures against buckling. It is also used to manufacture lightweight structures of metal honeycomb cores inside metal skins for the flight control component structures [29].
5.Delamination/Debonding Failure: Delamination is a failure mechanism in which lamina separate from each other in laminated composites. It occurs because of poor interlaminar fracture toughness and interlaminar stresses and results in loss in stiffness, strength, and expected life of the Page 14
material [7]. The simulation of delamination in composites is usually divided into delamination initiation and delamination propagation. Delamination initiation analyses are usually based on stresses and in delamination propagation analysis, energy release rate approach is used. The energy release rate can be evaluated using virtual crack closure technique (VCCT) which is based on Irwin’s assumption that when a crack extends by a small amount, the energy released in the process is equal to the work required to close the crack to its original length [10]. A well known example in which delamination failure resulted in the loss of an aircraft is the American Airlines Flight 587, in which the composite vertical stabilizer and rudder separated from the fuselage of the Airbus A300-600 aircraft, rendering the airplane uncontrollable [8]. Delamination growth can occur as a consequence of interlaminar stresses which can arise from fuel pressure variations, stiffness mismatch and in complex structures due to unanticipated loading such as excessive turbulence. It is important, therefore to improve the knowledge of delamination growth both theoretically and experimentally. It is also of interest in aircraft design to build up a database of material toughness on advanced carbon-fibre composites, (CFC), currently in use or considered for use in airframe structures [9].
6. Finite Element Analysis: Finite Element Analysis (FEA) was first developed in 1943 by R. Courant, who utilized the Ritz method of numerical analysis and minimization of variation calculus to obtain approximate solutions to vibration systems [11].Two types of analysis are generally used in the industry 2-D and 3-D modelling. While 2-D modelling conserves simplicity and allows the analysis to be run on a relatively normal computer, it tends to yield less Page 15
accurate results. 3-D modelling, however, produces more accurate results while sacrificing the ability to run on all but the fastest computers effectively [11].The finite element method using computers took off in the 1970’s when the Boeing Company launched a project to study stresses in the aircraft structure [1]. As the name suggests in finite element analysis the body being studied is divided into a number of small elements with their own physical properties such as thickness, coefficient of thermal expansion, density, Young's modulus, shear modulus etc. The connections points between these elements are known as nodes. The element’s geometry is dependent on the type of problem being studied. FEA saves a lot of money and time since new prototypes need not be made for the study, the design can be modified and studied on the computer itself. Smaller elements or a high element density is often used to improve the accuracy of the solution in regions where the stress gradients are high. During a finite element analysis study a balance between computer resources available and accuracy of results has to be achieved. The finite element analysis for this MSc thesis was carried out in ABAQUS software, which has been adopted by major corporations across all engineering disciplines as an integral part of their design process. ABAQUS offers a powerful and complete solution for simple to complex linear and nonlinear engineering problems, using the finite element method. In 2004 Abaqus was selected by Boeing to develop and market an add-on for the software, which incorporates the Virtual Crack Closure Technique (VCCT) proposed by Rybicki and Kanninen [6]. Perhaps the most important function of theoretical modelling is that of sharpening the designer's intuition; users of finite element codes should plan their
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strategy toward this end, supplementing the computer simulation with as much closed-form and experimental analysis as possible [12].
There are five basic steps involved in developing a finite element model of a physical system [13]:
Geometry definition
Discretization (i.e. meshing) of the geometry with a finite element mesh
Specification and assignment of material properties to finite elements
Specification of kinematic constraints
Specification of loading conditions
The different types of elements that may be used in FEA are as follows [13]:
Line Elements:
Line elements consist of 2 or more nodes that define the shape of a line. There are three distinct types of line elements: axial line elements have only stiffness in the axial direction, pure beam elements only have bending stiffness about one or more axes, and combined uniaxial /beam elements have both axial and bending stiffness’s.
a) Axial line elements: Also called uniaxial or spar elements are ideal for twoforce members which are common structural components in truss-type structures. Two-force members are mechanical components acted upon by two equal, opposite, and collinear forces.
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b) Bending line elements: Also known as pure beam elements, these elements are defined by 2 nodes. Each node of the beam element allows 3 degrees of freedom. Unlike uniaxial elements, beam elements do not allow stretching or compression stiffness. Only the component of force perpendicular to the elements axial direction is permitted, hence these elements are used for stress or strain calculations.
Surface or Area Elements
There are 2 types of area elements: 2 dimensional planar elements and 3 dimensional shell or plate elements:
a) Planar elements: These elements are used in 3 dimensional models which can be simplified, since it is cost effective to follow this approach. Planar elements permit only 2 degrees of freedom i.e. in X and Y direction. In solid mechanics there are three different types of planar analysis problems: 2D plane strain, 2D plane stress, and 2D axisymmetric. b) Shell or Plate Elements: Shell elements are used to model thin 3D structures usually acted upon by bending type loads. This element uses a different stiffness formulation than a standard solid element allowing higher accuracy.
Solid Elements: The simplest type of solid element is the linear tetrahedral with 4 nodes and the other two sold elements are the hexahedral or brick element with 8 nodes and the quadratic hexahedral element with 20 nodes.
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7. Structural Health Monitoring System
A better understanding of how delamination occurs in aircraft structures can lead to the development of real time structural health monitoring systems.
Figure 5: Equating SHM to the Nervous system
[14]
Structural health monitoring (SHM) can be imagined as the nervous system in the structure of an aircraft. Different types of sensor, some embedded in the airframe, detect cracks, corrosion, delamination and other damage and simplify their assessment [14]. A Real time structural health monitoring system will significantly reduce the risk of aircraft accidents due to structural failure and also reduce downtime of aircraft in maintenance hangars thus increasing profits for airline companies.
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Figure 6:SHM flow chart
[15]
Some structural health monitoring technologies which can be embedded into the aircraft structure are listed below: Comparative Vacuum Monitoring (CVM):
Figure 7: CVM sensor
[16]
CVM Sensor [17]
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CVM technology is based on the principle that a small volume maintained at a low vacuum is extremely sensitive to any ingress of air. When a crack develops it forms a leakage path between the atmospheric and vacuum galleries which can easily be detected. Self adhesive elastomeric sensors have been developed for this purpose. When a crack propagates from outside or from an atmospheric gallery, the seal between the atmospheric and vacuum galleries are broken which is detected by the CVM system [18]. Carbon Nanotube Network: With advancements in nano technology highly efficient sensors without any significant weight penalty can be developed, Carbon Nanotube are ideal sensors for incorporation into the structural health monitoring system of an aircraft. Carbon Nanotubes exhibit a behaviour called piezoresistivity i.e. change in resistance with strain. Such a sensor could measure large strain and form a grid over a large area of a structure for structural health monitoring (SHM) applications. Also, unlike other smart materials, CNTs are potentially simultaneously structural, functional and smart materials because of their load carrying capability, high thermal and electrical conductivity and sensing properties [30]. Acoustic Emission Sensor: Acoustic Emission (AE) is a phenomenon of all materials that when forces are applied stress waves are propagated through the material structure, which are measurable with suitable sensors. AE sensors are piezo-electric elements in most cases. They transform the stress waves into a voltage, which can be analysed with a suitable system. The frequency response of the sensors must be suitable for the frequency Page 21
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
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(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 23
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 24
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.
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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 26
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)
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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
Crew: 2
Max Cruising Speed: 470 Knots
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.
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(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.
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Fuel-Fraction in Cruise Segment: Flight Condition 1 Input Parameters
R
2500.0
nm
V
470.00
kts
cj
0.500
lb/hr lb
L/D
12.31
Output Parameter Mf f
0.8057
Advanced Aircraft Analysis 3.12 Project
09/14/09
2:16 PM
(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.
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(4) After the regression curve was defined the number of passengers and crew was entered into the program along with their estimated weights.
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(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 32
(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.
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Airplane Geometry W ing
Horiz ontal Tail
Ventral Fin
Fus elage
Vertical Tail
Canard
V-Tail
Angles
Landing Gear
Airplane
AeroPack
Scale
W ing Ty pe Selection Straight Tapered
Cranked W ing
Fuel Volume
Select Input Parameters Combination AR, S, , c/4 AR, S, c , r c/4
b, c , c , r t c/4
Adv anced Aircraft Analy s is 3.12 Project
09/14/09
Flap/Aileron/Tab
Chord Length
AR, S, , LE
(8)
3:43 PM
x mgc
cr
= 3.91 ft w
= 13.40 ft w
c w = 10.10 ft
ct
= 5.80 ft
w
y mgc
= 13.35 ft w
bw /2 = 30.75 ft
(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.
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(10)
Fuselage Geometry: Flight Condition 1 Input Parameters Xapex
Yapex
f
0.00
ft
Zapex
f
0.00
ft
if
f
0.00
ft
(X,Z)apex
0.00
deg
(X,Y,Z)f us
Nf
Apex is included
f
16
stations
Fuselage Coordinate System
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
fus
ft 1
y
fus
ft 1
z
fus
ft 1
y
fus
Advanced Aircraft Analysis 3.12 Project
ft 2
z
fus
ft 2
09/14/09
y
fus
ft
z
3
fus
ft 3
y
fus
ft z 12
fus
ft 12
fus
y 12
fus
ft z 23
fus
ft 23
fus
23
4:12 PM
(11)
Page 35
50.00
Afus
ft
30.00
Nose Center Tail
Area Afus
i
2
Z-location z cl
f
i
ft
z cl
f
0.00
20.00
-50.00
-100.00
10.00
-150.00
0.00
-200.00
-250.00 0.00
10.00
20.00
30.00
40.00
50.00
60.00
70.00
-10.00 80.00 90.00 100.00 Fuselage Station, x f /lf %
(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 aircraft to become unstable.
Page 36
x mgc = 0.08 ft h
c t = 5.29 ft h
c h = 7.07 ft c r = 8.59 ft h
y mgc = 4.87 ft h
bh/2 = 10.58 ft
(13)
x mgc = 0.61 ft v
c t = 7.93 ft v
c v = 10.80 ft c r = 13.22 ft v
z mgc = 2.73 ft v
bv = 5.95 ft
(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. Page 37
Nacelle Geometry: Flight Condition 1 Input Parameters Xnose
41.65
ft
Ynose
4.60
ft
Znose
1.0000
ft
n
n
n
0.00
deg
(X,Z)apex
n
0.0
deg
(X,Y,Z)n
n
0.0
deg
Nn
in
Apex is not included
n
Nacelle Coordinate System
7
stations
Output Parameter
Coordinates Defined
Nacelle 1 Table: double click for Cross-Section Dialog Station
x
n
ft 1
y
n
ft 1
z
n
ft 1
y
n
ft 2
z
n
ft 2
y
n
ft
z
3
n
ft 3
y
n
ft 12
z
n
ft 12
n
y 12
n
ft 23
z
n
ft 23
n
23
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
Advanced Aircraft Analysis 3.12 Project
09/22/09
4:18 PM
(15)
Page 38
Loads 3.00
Maneuver Diagram Gust Diagram
Load Factor n g
2.00
1.00
VS
VA
VB
eas
VC
eas
eas
VD
eas
0.00
-1.00
-2.00 0.00
100.00
200.00
300.00
400.00
500.00
600.00
700.00 800.00 Speed, Veas keas
900.00
(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 safety applied to a commercial aircraft is defined to be 1.5 by FAR §25.303. The
Page 39
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 40
W ing Lift Distribution: Flight Condition 1 Input Parameters
Altitude
T
U1
CL
w cln p.of f
Sw
ARw
6.41
o
deg F
w
0.43
o
kts
c/4 w
13.0
deg
@M=0 rw
6.3025
rad
@M=0 tw
6.3025
rad
34000
ft
0.0
470.00
cl
3.0000
590.40
cl
2
ft
rw M=0
2.0
deg
tw M=0
2.0
deg
w
9.03
%
w
9.00
%
(t/c)r
-1
(t/c)t
-1
cl
max
rw
cl
max
g w
tw
1.134
1.0
deg
1.189
Output Parameters M1
cl
0.812
cl
10.7857
rw
rad
-1
o
-1
tw
10.7857
rad
rw
1.7
deg
Advanced Aircraft Analysis 3.12 Project
09/18/09
o
a
tw
w
1.7
deg
1.0
deg
3:42 PM
(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.
Page 41
4.0000
Take-off Gear Down Take-off Gear Up Clean Landing Gear Down
Lift Coefficient CL
S = 590.40 ft
2
3.5000
3.0000
2.5000
2.0000
1.5000
1.0000 0.0000
0.2500
0.5000
0.7500
1.0000 Drag Coefficient, CD
1.2500
(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 landing gear up etc. L/D from W eights: Flight Condition 1 Input Parameters WTO
22571.5
Sw
lb
CD
2
590.40
ft
0
BDP
0.0161
clean,M
clean
0.0621
M ission Profile Table W
V kts
C L
C D
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
begin
lb
W
Advanced Aircraft Analysis 3.12 Project
F
lb
W
M ission Profile
begin
09/18/09
F
lb
h
ft
used
4:05 PM
(18)
Page 42
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. 1.00
(T/W)TO
Stall Speed Clean Stall Speed Take-off Distance TTO = 0 deg F
0.90
CL
0.80
Maximum Cruise Speed Landing Distance TL = 0 deg F
= 1.70
max
L
W TO = 22571.46 lb
0.70
CL
0.60
= 0.70
max
L
0.50
0.40
0.30
0.20 CL CL
max
0.10
CL
max
= 5.00
max
= 4.00
TO
TO
= 3.00 TO
0.00 0.00
25.00
50.00
75.00
100.00
125.00 (W/S)TO
150.00 lb 2
ft
(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.
Page 43
(20)
Page 44
2. Structural Analysis: The second phase of the project i.e. structural analysis and modelling was carried out in Abaqus 6.9.The primary focus of analysis was on the wing of the aircraft. The wing was cut in half for ease of calculation and modelling as the behaviour of the left and right wings would be similar. Modelling Structural Elements Wing/Skin: The wing was imported from Aeropack into Abaqus as a part; it was then cut along its mid span using the geometry repair function by removing the shell faces. This was done as only half of the wing was required for analysis and for easy insertion of structural components.
Page 45
I-Beam Flanges: To create I-beam flanges which followed the curvature of the wing, datum lines were made at locations where the spars needed to be inserted.
Partitions in the wing skin were made at these datum line locations and then a copy of the partitioned wing was made, in order to cut the unrequired faces of the skin to form the I-Beam flanges
Page 46
I-Beam/Spars Once the flanges were obtained, the beams had to be inserted into the flanges to obtain the I-beams which would make up the spars for the wing skin. In order to this the coordinates were obtained from the flanges and then sketched using the Abaqus Sketcher.
2-D side-view sketch of the beams were made and then extruded width wise by 0.025 meters. The beams were then assembled and merged with the flanges to form the IBeams.
Page 47
Ribs: The ribs were then sketched and extruded the same way as the spars. More coordinates were needed for the ribs as they followed the airfoil shape which was more complex than the spars.
Page 48
Page 49
Material Assignment Material properties now had to be assigned to all the parts, for this purpose materials were created in Abaqus by giving the materials mechanical properties such as Density, Young’s modulus and Poisson’s ratio .
Sections were then assigned to each face of the irregular parts separately, since in Abaqus only sections having the same geometry can be assigned material properties together.
Page 50
Mesh: After the section assignments were completed, the parts were meshed individually. The skin was assigned hexagonal elements and the solid parts were given quadratic elements. The spacing between elements was reduced at locations where it was believed that stresses would be higher in order to get an accurate picture while analyzing the assembly. A number of iterations had to be performed till the right mesh for analysis was obtained.
Page 51
Assembly:
After each part was meshed, the ribs were assembled into the spars and then the sparrib assembly was assembled into the wing. This was done by a series of rotations and translations which took some time to master as these manoeuvres had to be very accurate. The model was then ready for analysis.
Page 52
Analysis:
Vibrational Analysis:
In order to check and verify the behaviour of the model, a vibrational analysis was conducted for the first 10 natural frequencies.
Page 53
Steps
In the ‘steps’ module of Abaqus, the nature of analysis was defined. In this case the frequency step was chosen and the number of Eigen values was entered as 10 for the first 10 natural frequencies.
Constraints:
Since the model was made of different parts, constrains had to be assigned so that the model did not break apart during analysis. The flanges were constrained to the skin of the aircraft and the beams. The ribs were also constrained to the wing skin and spars. In Abaqus the inner or outer surface selections are determined by 2 colours i.e. Brown= outer surface and Purple= inner surface.
Page 54
Page 55
Boundary Conditions:
To simulate the wing being attached to fuselage, the wing skin root and the ends of the spars were encastered preventing rotation and translation in all directions at this location.
Interaction Properties:
In order to study debonding interaction properties could be assigned instead of boundary conditions and constraints. However due to the lack of time the interaction properties were not defined.
Page 56
Analysis:
The initial results were very disappointing with large perturbations in the wing skin and in one trial the I-beams broke apart and came out of the skin.
Page 57
The constraints were then adjusted and it was also found that the skin thickness had to be increased. The material properties were also researched again and adjusted as some errors had crept in during conversion from imperial to S.I. units.
The behaviour of the model for the first 10 natural frequencies was then successfully obtained with no perturbations in the skin:
Mode 1:
Page 58
Mode 2:
Mode 3:
Page 59
Mode 4:
Mode 5:
Page 60
Mode 6:
Mode 7:
Page 61
Mode 8:
Mode 9:
Page 62
Mode 10:
Page 63
Linear Elastic Analysis:
A linear elastic analysis of the wing was then carried out at steady level flight during cruise. Since at this condition lift is equal to weight and only half of the wing was being analyzed, the spars were assigned a load of half the weight of the aircraft and the skin was assigned an evenly distributed lift which was equal to half the aircraft weight, in the opposite direction. The weight of the aircraft at cruise was obtained from AAA.
Page 64
The linear elastic analysis job was then submitted yielding the following result. The undeformed shape is shown as shadow under the deformed shape.
Page 65
Conclusions: Results A conceptual design of an 8 seater business jet was completed. Vibrational and linear elastic analysis on its carbon fibre composited wing was also done. A procedure has now been established to design an aircraft in AAA and then design and analyze its structural components in Abaqus. Any changes that are required after structural analysis for example change in wing span; root/tip thickness etc. can easily be reinserted into AAA in order to analyze the effect of these changes on performance and if needed the design can be changed and revaluated in Abaqus till an optimum design is achieved.
Page 66
Further Work This project has great scope for further work. Due to time constraints only vibrational analysis and Linear Elastic analysis in steady level flight of the wing could be completed. If time permitted analysis of delamination could be performed and linear elastic analysis in other flight conditions could be performed on the wings. The fuselage and other components of the aircraft can also be given structural attributes and analyzed. Different materials could be assigned and changes produced could be studied. Debonding between skin and spars can be now studied, and delamination within the sandwich panes making the skin can be also analysed. Doing this project helps the student understand both the structural analysis and design process of aircraft design.
Project Management The literature review and learning how to use the Abaqus Software were done simultaneously. The aircraft design software AAA was taught as part of the course and therefore there was no need to learn it again for the dissertation. The aircraft design was done in the Howell building and structural analysis was done in the Michael Sterling building of Brunel University. There were some minor delays caused to the project due to upgrades done in the lab, however this was accounted for as a number of copies of the data were made.
Page 67
Gantt chart
Page 68
References: 1) Aircraft Structures for engineering students by T.H.G. Megson 2) Handbook of Adhesives and Sealants by Edward M. Petrie 3) Free online private pilot ground school (http://www.free-online-private-pilotground-school.com/aircraft-structure.html ) 4) Aircraft design a conceptual approach by Daniel P. Raymer 5) About.com(Composites/Plastics) (http://composite.about.com/cs/miscellaneousnews/a/bpr_abaqus.htm) 6) Delamination in Composite Materials Dr. Richard Chung
San Jose State
University 7) Materials Examination of the Vertical Stabilizer from American Airlines Flight 587 1National Transportation Safety Board, NASA Langley Research Centre, 8) A numerical and experimental investigation of delamination behaviour in the DCB specimen,Joakim Schöna, Tonny Nyman, 2002 9) Mixed-Mode Decohesion Finite Elements for the simulation of delamination in composite Materials P.P. Camanho ,C.G. Davila 10) Peter Widas ,Virginia Tech Material Science and Engineering 12) Finite Element Analysis, David Roylance , Department of Materials Science and Engineering, M.I.T. 13) Biomesh (www.biomesh.org) 14) European Aeronautic Defence and Space Company (EADS) (http://www.eads.com/800/en/madebyeads/endurance/shm.html) Page 69
15) Structural Health Monitoring for Life Management of Aircraft, Sridhar Krishnaswamy , North Western University ,U.S.A 16) CVM, Holger Speckman Airbus, Bremen, Germany 17) Sandia National Laboratories team leader Dennis Roach 18) Structural Health Monitoring ,Fu Ku Chang,2003 19) Composites and Advanced Materials ,U.S. Centennial of Flight Commission(http://www.centennialofflight.gov/essay/Evolution_of_Technolog y/composites/Tech40.htm) 20) Carbon fiber reinforced plastics in aircraft construction, C. Soutis, 2005 21) Star Telegram (www.star-telegram.com) 22) “Design considerations for composite fuselage structure of commercial transport aircraft”, G.W. Davis ,I.F. Sakata, NASA Contractor Report CR-159296 23) Online Video Lecture Series on Computational Methods in Design and Manufacturing by Dr. R. Krishnakumar, Department of Mechanical Engineering, IIT Madras. (http://nptel.iitm.ac.in/) 24) Aeronautics Learning Laboratory for science technology and research(http://www.allstar.fiu.edu/aero/flight12.htm) 25) “Multifunctional self healing and morphing composites”, T. Duenas*1, E. Bolanos2, E. Murphy2, A. Mal3, F. Wudl2, C. Schaffner2, Y. Wang3, H. T. Hahn3, T. K. Ooi4, A. Jha1,2007, US Army Aviation and Missile Research, Development, and Engineering Centre
Page 70
26) “Intelligent Material Systems Using Epoxy Particles to Repair Micro cracks and Delamination Damage in GFRP”, M. Zako and N. Takano,1999, Department of Manufacturing Science, Osaka University
27) “Self-healing polymer composites”, R S Trask, H R Williams and I P Bond, Department of Aerospace Engineering, University of Bristol 28) “Use of epoxy/multiwalled carbon nanotubes as adhesives to join graphite fibre reinforced polymer composites”, Kuang-Ting Hsiao, Justin Alms,Suresh G Advan,2000
29) UIUC Airfoil Coordinate database(http://www.ae.uiuc.edu/mselig/ads/coord_database.html) 30) “A carbon strain sensor for structural health monitoring”, Inpil Kang, Mark J Schulz
University of Cincinnati ,2006
31) “Intelligent Structural Health Monitoring (SHM) of Composite Aircraft structures using Acoustic Emission sensors”,Dirk Aljets ,2005 32) Net composites (www.netcomposites.com) 33) Azom materials (www.azom.com) 34) Composites and advanced materials (http://www.centennialofflight.gov/essay/Evolution_of_Technology/composites/Tech 40.htm)
35) “Wing instability of composite wing aircraft” Mahmood ,Fatholla,University of Iran
Page 71
36) “Flutter prediction, suppression and control in aircraft composite wings”,Nagarjuna, Cranfield university
Page 72
Appendix : Guidelines on interchanging between Abaqus and AAA 1) The design should be exported to Aeropack after it is completed in AAA 2) The file format used to export the model from Aeropack to Abaqus should be IGES 3) While exporting the model from Aerpack care must be taken about the units used as this can effect the whole project. Abaqus uses the same units the user has provided from the beginning and has no predefined units. 4) The whole aircraft can’t be meshed as a whole but can be meshed separately, however this provides little benefit to structural analysis as the model imported from Aeropack is a shell and structural components have to be added to it. 5) Ideally a single component from Aeropack should be imported and given structural attributes. 6) To get coordinates from the wing skin in order to model structures like spars and ribs, the skin or flanges should be given an arbitrary mesh and then the coordinates of the nodes can easily be found using the query option in the tools menu.
Page 73
Email from Grob Aircraft Dear Mr. Narayan,
All Grob aircraft are produced by use of wet lay-up composite materials. The resin system for motorplanes is L20/SL (today called ERP L20 / EPH 960). Some gliders are produced from the Scheuffler resin system L285 / H285, H286, H287. The very old gliders from Epicote / Laromin.
Fibre Fabrics are: Interglas 92110, 92125, 92140, 92145, 92146 and comparable fabrics. Carbon fabrics: Mainly 98141 or ECC 452, also ECC459. Glas Fibre Rovings: Vetrotex EC9, Carbon HTA Rovings.
We hope that helps.
Best regards Jörg Unbehend
Joerg Unbehend Head of Design
*** Bitte beachten Sie meine neue E-Mail Adresse ***
Phone: +49 (0) 8268 998 424 Fax:
+49 (0) 8268 998 221
GROB AIRCRAFT AG Lettenbachstrasse 9
Page 74
86874 Tussenhausen-Mattsies Germany
www.grob-aircraft.com
Sitz Tussenhausen-Mattsies, Amtsgericht Memmingen, HRB 13686
Vertretungsberechtigter Vorstand: Johann Heitzmann, Andre Hiebeler, Andreas Konle
Aufsichtsratsvorsitzende: Antoinette Hiebeler-Hasner
-----Ursprüngliche Nachricht----Von: Gehling Ulrich Gesendet: Montag, 29. Juni 2009 09:15 An: Unbehend Joerg Betreff: WG: Grob-Contactform Message (Product Support)
Bitte um KURZE Antwort direkt an Sender: Mr Anirudh Narayan Danke
-----Ursprüngliche Nachricht----Von: Vodermeier Rudolf Gesendet: Montag, 29. Juni 2009 07:59 An: .GF Betreff: WG: Grob-Contactform Message (Product Support)
Page 75
-----Ursprüngliche Nachricht----Von:
[email protected] [mailto:
[email protected]] Gesendet: Sonntag, 28. Juni 2009 17:01 An: -EVL-Productsupport Betreff: Grob-Contactform Message (Product Support)
Page 76
