Q

Assignments is about the challenges of current world are constantly fueling the discovery and development of new kinds

Home, - AIRCRAFT WINGS AND FUSELAGE

INTRODUCTION

The airframe of a fixed-wing aircraft is generally considered to consist of five principal units, the fuselage, wings, stabilizers, flight control surfaces, and landing gear. Helicopter airframe consist of fuselage, main rotor and related gearbox, tail rotor and the landing gear. The airframe components are constructed from a wide variety of materials and are joined by rivets, bolts, screws, and welding or adhesives. The aircraft components are composed of various parts called structural members (i.e. Stringers, longerons, ribs, bulkheads, etc.). Aircraft structural members are designed to carry a load or to resist stress. A single member of the structure may be subjected to a combination of stresses. In most cases the structural members are designed to carry loads rather than side; that is, to be subjected to tension or compression rather than bending. Strength may be the principal requirement in certain structures, while others need entirely different qualities. For example, cowling, fairing, and similar parts usually are not required to carry the stresses imposed by flight or the landingloads. However, these parts must have such properties as neat appearance and streamlined shapes.
OBJECTIVES
1. To find out the normal uses of materials used in the manufacture of aircraft wings and fuselage.
2. To describe the history and development of engineering materials used in the manufacture of the aircraft structures and fuselage
3. To understand the major property requirements for instance the specifications in terms of range of values for a material to be considered for selection
4. To point out the unique properties of the engineering materials used in manufacturing of aircraft wings and fuselage.
5.To find out the technical ,economic ,legal ,sustainability and social hurdles that could hinder the application of engineering materials as used in the aircraft wings and fuselage

 


BACKGROUND INFORMATION
History and development of engineering materials used in manufacturing of aircraft wings and fuselage.
THE EVOLUTION
Since the beginnings of aircraft history, a lightweight structure strength was one of biggest challenge for pioneers together with the aerodynamic of lifting surfaces, an appropriate thrust source and accurate controls. The first controlled successful flight was not only an fortunately attempt, being the result of a scientific approach of Wright brothers, based on a carefully study of the research of their forerunners like Sir George Cayley, Otto Lilienthal or Octave Chanute Being the author of an well-known monography, Octave Chanute, an successful civil engineer (bridges and railroads) acted as consultant for Wright brothers. It was Chanute's idea to combine the biplane (intensively tested by Lilienthal) with the latticed beam, resulting a high inertia momentum. Lilienthal bird had wing biplane with upper plane mounted on a mast; having wooden ribs, the margins of the wing were wire attached by mast. Wright's Flyer had the two wings of the biplane acting as two flanges of a Pratt beam; having three bays each side, those were reinforced with diagonal wire bracing, this being a lightweight innovation in order to replace the diagonal bracing. Latticed beams of early aircrafts were composed by longerons, columns and diagonal members (wire bracings). The most used section was square but there were also exceptions The longerons were initially made from ash or hickory, spruce being used later for weight saving. Because of problems related to wooden/ fabric construction, manufacturers searched for solutions to eliminate the lack of durability, crashworthy and damage tolerance, flammability, anisotropy, wire bracing and so on. Between latticed beams and semimonocoque, there were a lot of concepts as:
Beam without diagonals, with external stiffening skin;
Beam with longerons, frames column members replaced by frames with stiffening skin;
Beam with a system of stringers replacing the longerons, reinforced by circular soft frames and
swept skin;
? Beams with frames, diagonal wire bracing and skin
? As materials there were employed:
? Wood, aluminum or steel tubes for longerons, columns, frames and diagonal members
? Plywood, veneer, steel or aluminium sheet metal (plain or corrugated) for skin and frames
? Aluminum or steel for connections and fittings.
It can be concluded a big mixture of members type and materials was used; in a logical approach these concepts leads to the semimonocoque, but many of this solutions were used many years after successful semimonocoque plane. Even semimonocoque appeared in 1912 showing improved features and characteristics, because of manufacturing costs it was not implemented in the design of the new aircrafts of the era. One of the biggest requirement of the First World War was the short development time, as a response to battlefield request. There were aircrafts which were released only after 3 - 4 months after first hand sketches. The transition from bamboo and wood to metal led to the requirement of new joining techniques.
Starting from 1907, Anthony Fokker used on his airplanes welded structure, joining up to eight members also with hinged struts and wires Since 1919, Camm stated that metal tubing is the most practical form in which steel can be used on aircrafts . In is presented a joint with gusset like curved tublets for wire connections. There were used also spruce filled tubes to prevent local buckling. Welded tubes structures were known and appreciated for accuracy and productivity since First World War, but welding techniques needed more progress in order to compete with wooden structures. Flight magazine noted in 1918, that Fokker structure welded nodes are the result of an "excellent workmanship". Early problems with welding led to the idea the welding depends a lot of the welder skills, this being propagated until present time
Latticed beams of early aircrafts were composed by longerons, columns and diagonal members (wire bracings). The most used section was square but there were also exceptions. The longerons were initially made from ash or hickory, spruce being used later for weight saving Wire bracing even having low weight, needed a skilled mechanic to adjust and maintain them, having the risk of structure failure if one wire fails. For these reasons, manufactures tend to replace wires with diagonal members, Camm in 1919 noted that the design and type of fittings employed for connecting the latticed beam members varies greatly, being one of the distinctive constructional details of a plane, being mainly the result of desire for originality of each individual designer and had to disappear with the progress of the industry .Even this aspect leads to manufacturing and productivity problems, it was propagated until nowadays. In the early days of aviation, the fuselage fittings were made of aluminum alloy, but after in 1915 the standard fittings were from stamped steel. Because of problems related to wooden/ fabric construction, manufacturers searched for solutions toeliminate the lack of durability, crashworthy and damage tolerance, flammability, anisotropy, wire bracing andso on. Between latticed beams and semimonocoque, there were a lot of concepts as:
? Beam without diagonals, with external stiffening skin;
? Beam with longerons, frames column members replaced by frames with stiffening skin;
? Beam with a system of stringers replacing the longerons, reinforced by circular soft frames and swept skin;
? Beams with frames, diagonal wire bracing and skin
? As materials there were employed:
? Wood, aluminum or steel tubes for longerons, columns, frames and diagonal members
? Plywood, veneer, steel or aluminum sheet metal (plain or corrugated) for skin and frames
Aluminum or steel for connections and fittings
It can be concluded a big mixture of members type and materials was used; in a logical approach these
concepts leads to the semimonocoque, but many of this solutions were used many years after successful semimonocoque plane. Even semimonocoque appeared in 1912 showing improved features and characteristics, because of manufacturing costs it was not implemented in the design of the new aircrafts .The first reliable structure for aircraft was the latticed beam, solution inspired from civil engineering used for fuselage and wing, it was a standard up to thirties, until semimonocoque (originated from shipsengineering) finally demonstrated its superiority. Biplane wing was used until monoplanedemonstrates its aerodynamic superiority and its internal structure reached a level where the beam made by twoplanes was no needed anymore.Latticed beams were made by wooden members with wire bracing until steel or aluminum tubesreplaced them, needing a more simple joining. The transition from latticed beam fuselage to semimonocoquewas somehow superposed over the transition from wood to metal, but they were almost independent processeseven they interfered a lot. Welding is used from the early aircraft but in that time the technology was not mature. For this reason, for more than 30 years manufactures searched for alternative to welding, joining type leading even to members different construction. Latticed beams are used only in limited application for fuselage, having welded nodes. Weld is used in many other application on aircrafts but not in latticed beam structures (landing gears, empennage, seats, etc.).
FUSELAGE

Fuselage is the elongated structure, of approximatelyStreamline form, to which areattached thewings andTail unit of an airplane. In general, it is designed tohold the passengers.
Fuselage, length ofthe distance from the nose of the fuselage (Including the engine bed and radiator, if present) to the after end of the fuselage, not including the controland stabilizing surfaces.

 

Normal uses materials used in the aircraft wings and fuselage

FIGURE 1(aircraft parts)
The structure of an airplane is referred to as the airframe and is composed of five major units including wings, fuselage, stabilizingsurfaces (fin and tail plane), landing gear and the flying controls surfaces as shown in figure above.
For the aircraft structures to operate normally it must be strong enough to withstand the forces acting upon it during the stages of flight without failure and distortion. Furthermore they must be joined together by means of screws, bolts, rivets, welding etcetera. Whatever methodis used to join them it must be strong enough to withstand the loads to which the junctions are subjected to.
2. The wings support the aircraft in flight and thereforeit must be made of materials that are strong enough to withstand the aerodynamic forces without bending excessivelyor twisting.However has this forces vary with different flight speeds or during turbulence the wings must have the ability to flex.The same applies with the junction between the wings and fuselage
3. When the rudder and elevators are used the forces acting upon them tend to twist and bend the fuselage which must be strong enough to resist this. It is also important that whilst the wings can be to flex up and down do not twist when the ailerons are used.
4. When the elevators are deflected up and down there is a twisting force (torque) applied to the horizontal stabilizer and its attachment to the fuselage. Both must be strong enough to resist the twisting force but the stabilizer must be supple enough to flex, or bend otherwise it might snap. The same requirements exist for the fin when the rudder is deflected right or left.
5. The landing gear must be strong enough not only to support the weight of aircraft on the ground but to withstand the shock of the landing, the twisting loads when the aircraft turns during taxiing and the bending loads during touchdown. All this applies to the points of attachment of the landing gear and the airframe.
This therefore points out that the designer must consider the factors above to come up with an aircraft made up of materials that are strong enough to withstand the loads to which the airframe is subjected to. The materials must be however flexible where necessary to absorb the changing loads which on the other hand could be rigid to prevent twisting
Fuselage construction.
Besides providing accommodation to the passengers, freight, system,instruments, andcrew, the fuselage must be able to withstand the stresses of flight, they are typically the torsion from the empennage (rudder and elevators) and the propellers from the single engine aircraft, bending on the touchdown and tension and compression from the wings inflight. There are three main forms of fuselage construction known as steel tube (truss),monocogue and semimonocoque.
1. Truss type of fuselage construction.
The truss type of fuselage comprises of framework made, in modern aircraft, of steel tubes. The principle components are longitudinal tubes called longerons, with intermediate diagonal braces.
The basic concept of truss construction is that the compression and the tension stresses due to bending that a fuselage is subjected to are alternately carried by the truss components. When the bending loads are reversed the loading of the truss members is reversed and so the stresses are spread evenly over the whole structure and therefore avoiding the concentration at one point. Although steel tubing's is the material used commonly in truss construction nowdays,wood and aluminum has been extensively used in the past often with steel wire forming some of the bracings members. Truss construction of fuselage is only limited to aircraft fuselages. The fuselage skin is made up of thin gauge aluminumsince it carries no load. In earlier aircraft it was often made up of fabric and wood.

FIGURE 2(type of truss construction)
2.monocoque type of fuselage.
This name means single shell and in this type of construction the strength to maintain the fuselage rigidity and withstand stress is all in the fuselage skin. There are no bracing members as compared to the truss type of construction but only formers to maintain the desired shape of the fuselage. Since the skin must take all the loads this type of construction is unsuited to large diameter fuselages because the skin thickness necessary would incur a high weight penalty. Hence the monocoque type of construction is limited to small and narrow fuselages. The material commonly used for monocoque construction is high strength aluminumalloy, 2024 duralumin being a typical example.

FIGURE 3(monocoque type of fuselage)
3. Semi monocoque
Neither truss nor monocoque construction is suitable for most aircraft fuselages especially where large and pressurized aircraft are concerned. Because of this a form of semi-monocoque construction is used which employs longerons to brace the load bearing skin material and take some of the loads.
Shorter longitudinal members called stringers supplement the longerons.Formers called frames, rings and bulkheads maintain fuselage shape.The main advantage of this form of construction is that it is capable of maintaining its structuralintegrity even in the event of considerable damage, since the loads and stresses are spread all over the whole structure rather than being concentrated on the frames or skin.
The longerons and stringers absorb the tensile and compression stress due to bending whilst torsional stress is taken up by the skin. The longerons and stringers are also the attachment points for the skin.
The materials used in semi-monocoque construction are principally metal, with high strength aluminum alloy being the commonest, especially, in smaller aircraft. In larger aircraft steel and titanium alloys are often used for major load bearing components.Secondary and non-load bearing components are increasingly made from fiberglass, Kevlar, graphite based compounds and composites.Cabin floors are often made from aluminum and fiber glass honeycomb sandwiched with aluminumsheeting. In many aircraft fuselages, especially smaller types, combination of structural methods may be used. Some cesna designs, forexample, use steel truss construction for the forward fuselage and cockpit area and semi-monocoque for the rear fuselage and tail cone.
Large transport aircraft fuselages are usually of semi monocoque construction and is formed of a number of sections joined end to end. The simplest comprises ofa streamlined nose section including the flight deck, a parallel sided cylindrical cabin section to which the wings are attached and tapered tail section carrying the empennage. Cabin windows are usually made of strong plastic material such as Perspex, and this is also used for the cockpit windows of un-pressurizedaircraft. The cockpit windows of a pressurized aircraft are usually made from strengthened glass.


FIGURE 3(semi-monocoque construction of fuselage)
In designing an aircraft, every square inch of wing and fuselage, every rib, spar, and even each metal fitting must be considered in relation to the physical characteristics of the metal of which it is made. Every part of the aircraft must be planned to carry the load to be imposed upon it. The determination of such loads is called stress analysis. Although planning the design is not the function of the aviation mechanic, it is, nevertheless, important to understand and appreciate the stresses involved in order to avoid changes in the original design through improper repairs.

 

Wings construction
The wings are generate virtually all of the lift that keeps the aircraft airborne. The wings therefore support the remainder of the aircraft. Thus in flight there is a considerable upward bending force acting upon the wings and is largely concentrated at the point of attachment to the fuselage. In addition the ailerons when deflected apply a twisting force about the lateral centerline of the wings. Consequently the wing structure must be strong enough to withstand the bending and torsional stresses, which are trying to deform the wing. The fuselage attachment points must be able to withstand the stresses imposed by the upward bending forces acting on the wings and by the twisting forces applied on the ailerons, both of which are trying to separate the wings from the fuselage
In some aircraft where the wings are necessarily of light construction the loads are in part taken by bracing struts and wires. In most cases the wings are designed in what is called cantilever principle, where structural rigidity is provided entirely by the wing structural members.
The bending stresses to which the wings is subjected may be carried by one or more transverse beams known as spars, or by building the wing into a box structure in which almost all the stresses are carried by the external skin. The latter is called the stressed-skin construction.
Torsional stress due to largely the effects of movement of the center of pressure is taken up by the chord wise ribs that give greater rigidity. The ribs also provide the aero foilshape. Stringers ran span wise between the spars to provide the attachment points for the skin and provide additional span wise rigidity
Wings of spar construction are either monospar having a single spar has the name suggests, two spar or multispar.

The wing spars of modern aircraft are made of metal formed into beam either by extrusion or built up construction. Some of the examples of spar beam constructions are

The material most commonly used is high strength aluminumalloy, although a few highly stressed aircraft, particularly military types, using titanium spar webs. The attachment points between the wings par and fuselage center section are often of titanium or steel alloy in large aircraft.
The wings ribs are the formers that maintain the aero foil section of the wing and have to be strong enough to resist the torsional stress tending to twist the wing. This forces are much less than the bending loads carried by the spars and the ribs are consequently of light construction. They are typically either pressed on one piece or built up on light aluminum alloy sheet.
Span wise stiffness is complemented by use of stringers running parallel to the spars to which the wing skin is attached. In spars constructed wings this skin is usually of light aluminum alloy sheet.
Stressed skin wings have no spars as such, but shear webs running span wise to which the ribs are attached. In turn the stringers are attached to the ribs and the stressed skin metal is riveted to the stringers to form a load bearing box.The stressed skin material is usually of high tensile aluminium the thickness ranging from about half a millimeter in small aircraft to as much as 16mm inlarge transport aircraft.
The bending stresses acting upon the wing can be alleviated to some extent by applying downward forces to oppose the upward forces of lift. This can be achieved by wing mounted engines and by the weight of the fuel in outboard fuel tanks. In aircraft with fuselage mounted engines, or compensate when the fuel in the outboard tanks has been used up, some aircraft are fitted with ailerons biased towards the up position to provide the stress relieving downward force at the outer wings.The wings of some aircraft are of cantilever design; that is, they are built so that no external bracing is needed. The skin is part of the wing structure and carries part of the wing stresses. Other aircraft wings use external bracing (struts, wires, etc.) to assist in supporting the wing and carrying the aerodynamic and landing loads. Both aluminum alloy and magnesium alloy are used in wing construction. The internal structure is made up of spars and stringers running span wise, and ribs and formers running chord wise (leading edge to trailing edge). The spars are the principal structural members of the wing. The skin is attached to the internal members and may carry part of the wing stresses. During flight, applied loads which are imposed on the wing structure are primarily on the skin. From the skin they are transmitted


Normal operating conditions of the engineering materials
An aero plane's airframe needs to satisfy a few conditions for it to be accepted into official service, with the most demanding ones being toughness and lightness. Besides that, it also needs to be cheaply accessed in an industrial level as well as easily tooled to keep the price down.
Aluminum
It's tough, it's light, it's plentiful, and an entire rolled sheet can be bought for less than you can buy steel. It's the perfect material.
But not all types of Aluminum can be used to make an entire aero plane out of, you would need a very particular alloy of Aluminum. Different alloys of Aluminum can have vastly different mechanical properties such as tensile strength, elasticity, yield strength,
5356 Al is used as a welding filler, not a designer material, although it has high strength, it is very prone to cracking and would fare horribly against the stress an aero plane would face.
Aerospace-gradeAluminum, are with good mechanical properties, particularly 7075 or 6061 Al. They've been utilized ubiquitously since the 50s-60s, and continued to do so now.But right now we're looking at a shift from Aluminum sheets to composite materials, with the first in line to do so being the Boeing 787, which is composed some 50% by a single composite - a mixture of Aluminum, Titanium, steel, and some other metals. This gives the aircraft much morestrength, reduce more weight, and improve maintenance than an atypical Aluminum skin in the past. Inmost of aircrafts operating today, they're all made from aluminum alloys. But each of them have different compositions of alloy. Most of composition used is 2 series and 7 series. 7 series are usually used in a very major parts or parts that receive so high load. The most popular compositions are Al-2024 and Al-7075. In aircrafts with special needs, they might replace aluminum with titanium, and it's alloy, most of the skin and stringers will be replaced with sandwich structure. It's a structure that arranged as skin-honeycomb-skin. In this case, the aluminum alloy would be replaced A large majority of airplane wings are made from sheet aluminum arranged to form a series of structural boxes. This arrangement was invented by Howard Hughes. In many larger airplanes, the boxes and the external skin form the fuel tank also. A strong aluminum alloy core (often 2024) with a pure aluminum outer sheet top and bottom. The alloy core is very strong, but subject to corrosion. The aluminum has little strength, but is very resistant to corrosion. Early airplane winds were built almost entirely of wood, with some steel cable to provide diagonal bracing. Then the skin of the wing was cotton fabric, or later Dacron fabric, stiffened with dope. I think a few aircraft are still made of very similar construction, but using aluminum or fiberglass structure instead of wood. With carbon fiber reinforced plastic...Many newer (usually smaller) aircraft have the wing (and sometimes the fuselage as well) made of composite plastics. The early examples were made strictly of fiberglass cloth and then either polyester or epoxy resin. A minor "problem" here is that fiberglass fiber is somewhat flexible, and allows the wings to bend quite a lot. Newer composite wings frequently use carbon fiber, especially for the wing spar. This lightens the wing a bit, but adds a good amount of stiffness.There are a few aircraft built with plywood wing structures and skin. These aircraft can never be left out in the weather.For all of the mainstream airliners, until very recently the answer was predominantly aluminum. An inside structure, either with box-like sections or a skeleton of ribs, spars and stringers over which the skin is fastened. The skin is a stressed part of the structure

The recent change has been to use a lot more composite materials. This has come incrementally, from using a few key components e.g. the Centre wing box which holds the two wings together on the Airbus A380, to far more extensive use of composites on Boeing 787 and Airbus A350. The exact composites are still variable, and given their relative infancy will continue to change, but mostly carbon fiber reinforced plastic (CFRP) for now.
In general, composite materials don't like impact damage, so it is expectected to see things like the leading edge of wings using a metal front piece to meet impact damage requirements, e.g. bird strike. Of recent, the airliner fleet in world-wide use has a mix of ages of aircraft. A very large number still predominantly aluminum structure, but the percentage of aircraft with significant amounts of composite materials is growing, and will continue to do so. One of the most important objectives for airplane designers is reduction of weight - build something that meets all the functional requirements while weighing as little as possible (within reasonable cost parameters, of course. Less weight means less engine thrust is needed, which means less fuel is burned. Less fuel burned means the airline makes more money. Airlines love to make money
Large commercial transport aircraft have traditionally been made with aluminum skin over analuminum frame. Aluminum is a great metal for airplanes since it's light and can be pretty strong when alloyed. But it has some issues over time, including corrosion and metal fatigue. With the introduction of the Boeing 787 a few years ago, large-scale use of carbon fiber reinforced plastic, aka composite, is proving itself as a lighter weight, corrosion-proof, and stronger replacement for aluminum skin. The Airbus A350 entering service later this year also is skinned with composite material, but uses panels over a framerather than one piece composite fuselage barrels like the 787.
Different aircraft require different building materials. Aircraft can be constructed from wood, fabric, many types of metal, or even composite materials (e.g. carbon-fiber, fiberglass). Early aircraft such as the Wright Flyer were built with wood and fabric. The frame of the Wright Flyer was made from spruce and ash and many surfaces were covered with muslin, a fabric. Most airplanes today are made out of aluminum, a strong, yet lightweight metal. The Ford Tri-Motor, the first passenger plane from 1928, was made out of aluminum. The modern Boeing 747 is an aluminum airplane as well. Other metals, such as steel and titanium, are sometimes used to build aircraft. Steel is heavy though, so not too much is used. Titanium is almost as strong as steel, has a medium weight, is heat resistant, and is corrosion resistant. The Lockheed SR-71 Blackbird, the world's fastest jet-propelled aircraft, is made of titanium. Composite materials such as graphite-epoxy are strong, but can weigh half as much as aluminum. These lightweight, customizable materials are becoming more popular. More than half of the materials used to make the Boeing 787 Dreamliner are composites. An Airbus jetliner is the product of highly-efficient cooperation across the company's global supply and manufacturing chains, as well as its decades of innovation for the air transport sector. In order to ensure the most efficient flight patterns possible, you must keep your aircraft clean and free of debris. The buildup can affect flight speed and performance. This includes bug and pest stain, sap, and oil that can collect around the wings or throughout the lower exterior of the aircraft. For cleaning, it's best to avoid the most corrosive cleaning products. These include most of the cleaners meant for household use. These may be great for kitchen counters and bathroom floors but can be detrimental to aluminum surfaces and damage sealants on aircraft, so definitely use caution when selecting them. The cleaning period is also prime time for inspecting your aircraft. A full inspection accompanying each cleaning could prevent subsequent maintenance issues or enable owners to catch developing issues before they get out of hand. Every aircraft owner has their own personal technique for polishing, but patience is the most universal asset to have when facing any polishing job. To maximize efficiency, it's a good idea keep several towels and extra supplies within reach for added convenience.

Back in the early days of aviation, wings were made of everything from ash wood for the frame members and double layered fabric would be used for the surface of the wings. The fabric was applied on the bias (45 degree angle) to increase the stiffness of the wings. [Wright Brothers, 1903]
Fast forward to the present day, and the wings are made of aluminum, with molded members and components and sheet aluminum riveted to make the upper and lower wing surfaces. Control surfaces, such as flaps, slats, spoilers (airbrakes), and ailerons, are also made of aluminum, but are controlled separately and independently.The greatest variety of materials can be found among recreational aircraft. Almost any combination of wood, aluminum, chromolly steel and composite (fiberglass or carbon fiber) for the aircraft structure to natural fabric, Dacron and other synthetics, plywood, aluminum sheet and composite for the aircraft skin. Composite aircraft may also use foam as a filler to give shape or sandwiched between layers of fiberglass for strengthening. Little or notitanium.

Specifications of aircraft wings and fuselage
Compliance statements are shown below:

1 Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
Calculated loads are limit loads.


2 Unless otherwise provided, the air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions.
Calculations are based on equilibrium of air and inertia forces.
Mass distribution is taken into consideration.


3 The simplified structural designcontains acceptable methods of analysis that may be used for compliance with the loading requirements for the wings and fuselage.
The simplified design criteria is not used.


4 Flight load factors, n, represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the airplane) to the weight of the airplane. A positive flight load factor is one in which the aerodynamic force acts upward, with respect to the airplane.
The requirement has been met.


5 Compliance with the flight load requirements of this section must be shown at each practicable combination of weight and disposable load within the operating limitations specified.
Every combination of weight and centre of gravity was considered.



7 Compliance with the strength requirements of this subpart must be shown at any combination of airspeed and load factor on and within the boundaries of a flight envelope that represents the envelope of the flight loading conditions specified by themanoeuvring and gust criteria
Every combination of airspeed and load factor on and within the boundaries of the flight envelope was considered.



8 The airplane is assumed to be subjected to the unsymmetrical flight condition Unbalanced aerodynamic moments about the centre of gravity must be reacted in a rational or conservative manner considering the principal masses furnishing the reacting inertia forces.


9 The airplane shall be designed for the loads resulting from the roll control deflections and speeds specified in combination with a load factor of at least two thirds of the positive manoeuvring load factor prescribed The rolling accelerations may be obtained by the methods given. The effect of the roll control displacement on the wing torsion may be accounted for by the givenmethod.


The airplane must be designed for the yawing loads resulting from the vertical surface loads specified

 



Loadings
As a fully aerobatic aircraft swift VLA is required to withstand +/-9G limit loads. Forinitial airframe stressing there is an ultimate+/-9 Initial mass estimation is then established
Both limit shear force and bending moment diagrams are then obtained and initial sizing of skins and frames could begin.A structure which is subjected to continual reversals of loading will fail at a load of less than would be the case for a steadily applied load. The failing load will depend on the number of reversals experienced. It can be seen in the example below that if the applied stress was 80% of the ultimate stress, the specimen could expect to fail after 100 applications but if the applied stress was reduced to 20% the failure would not occur until 10
Having calculated the maximum anticipated loads the designer arrives at a maximum compromise which gives a sufficient structural strength but keeps the airframe weight to minimum. This normally ensures that each of the various parts of the structure is designed to fail at an ultimate load of one and half times greater than the maximum applied load. The ratio of the ultimate load to maximum applied load is (1:5:1) which is the safety factor.
Million applications. A pneumatic system is fitted in most modern aircraft to supply some or all of the following
? Aircraft systems.
? Air conditioning
? Pressurization
? Aero foil and engine anti-icing
? Air turbine motors
? Engine starting
? Hydraulic power
? Thrust reverse
? Leading and trailing edge flap/slat operation
? Pneumatic rams, e.g. thrust reverser actuation
? Hydraulic reservoir and potable water tank pressurization
? Cargo compartment heating
Most of these systems use high volume low pressure airflow bled from the compressor stagesof a gas turbine engine,
Damage Tolerant Structure
Fail safe structures are rather heavy due to the extra structural members required to protect the integrity of the structure. Damage tolerant structure eliminates the extra structural members by spreading the loading of a particular structure over a larger area. This means that the structure is designed so that damage can be detected during the normal inspection cycles before a failure occurs.

The aircraft manufacturer will attempt to design an aircraft to take into account all the loadsthat it may experience in flight. There are various guidelines, formulae and experience to guidethem in the design of a good fail safe/damage tolerant structure.
Safe Life
The safe life of an aircraft structure is defined as the minimum life during which it is known thatno catastrophic damage will occur. Life-counts for components of assemblies may be recordedas number of flying hours, cycles of landing or pressurization events or even on a calendarbasis. After the elapsed life-count or fatigue cycle (typically pressurizations or landings hasbeen reached, the item is replaced or overhauled. In the interim (operational life) of the Aircraft,and to minimize the chances of failure due to fatigue, aircraft designers apply the principle ofFail-safeconstruction or Damage tolerance.
Fail Safe or Damage Tolerant Structure
Large modern Aircraft are designed with aFail-safeor Damage-tolerantstructure. This can be described as a structure in which a failure of a particular part is compensated for by an alternative load-path provided by an adjacent part that is able to carry the loads for a limited time period.Typically this is a structure which, after any single failure or crack in any one structural member can safely carry the normal operating loads until the next periodic inspection. True dualling of load-paths in common practice could be found in wing attachments and also in vertical stabilizer and horizontal stabilizer attachment points. Detection of faults is reliant upon a planned inspection program capable of finding such failures. In order to gain access to the vulnerable areas a certain amount of dismantling is necessary although the use of non-destructive testing (NDT) may be employed in less critical areas. The disadvantage of true dualling of load-paths is that it is fundamentally very heavy. Modern concepts of construction employ the‘semi-Monococque' style of construction where each piece of the Aircraft has its part to play in spreading loads throughout the Airframe and is tolerant to certain amount of damage. The programmed inspection cycle periodicy is determined on the basis that if a crack of detectable length has been missed at the first inspection, the structure will allow this crack to develop until a subsequent inspection before it becomes critical. The criteria of inspection cycles, Design Limit Loads, and Design Ultimate Loads are agreed at the time of certification.

Major property requirements (specifications)

Characteristics of the engineering materials used in the manufacture of aircraft wings and fuselage
Carbon Fiber-Reinforced Polymer (CFRP) Composites
Carbon is a high-performance fiber material that is the most commonlyin used reinforcementin advanced (i.e., non-fiberglass) polymer-matrix composites .The reasonsfor this are as follows:
1. Carbon fibers have the highest specific modulus and specific strength of all reinforcing fiber materials.
2. They retain their high tensile modulus and high strength at elevated temperatures;high-temperature oxidation, however, may be a problem.
3. At room temperature, carbon fibers are not affected by moisture or a wide variety of solvents, acids, and bases.
4. These fibers exhibit a diversity of physical and mechanical characteristics,allowing composites incorporating these fibers to have specific engineeredproperties.
5. Fiber and composite manufacturing processes have been developed that are relatively inexpensive and cost effective.

Technical, economic, legal, sustainability and social hurdles that could hinder the application of the engineering materials.
One of the difficulties in managing aircraft fleets is tracking data from past aircraft inspections torefine the assessment of the current condition of each aircraft or the entire fleet. A good knowledge of the current state of a component or aircraft is important for accurately determining the risk of failure. Electronic databases make storing the data easier. The challenge is getting the data into the database.
Challenges
The technical (e.g. recyclability), economic, legal, sustainability, and social hurdles that could hinder the application of the candidate engineering materials as used in aircraft wings and fuselage
Challenges of using composites in aircraft wings and fuselage
1. Orthographic nature of composite materials
2. Generation of design allowances for advanced composites still proves to be costly and time consuming
3. Extensive coupon test plans required to satisfy regulations.
3. Manufacturers are unwilling to share design allowances generated
4. Leads to consecutive test programs and design allowances being generated by multiple manufactures at a great cost each.
Methods of preventing aging of aircraft wings and fuselage
Penetrant testing
Penetrant method (penetrant testing) is used to reveal discontinuities opened towards the surface of the parts made of material that is not porous. The method depends on the ability of liquid to penetrate the discontinuity of the material on which it is applied. Ordinary penetration method is used for the detection of small cracks or gaps that go up to the surface and that may not be visible under normal visual inspection. Penetration methods can be used in most parts of thestructures that are available for its application. The process of penetrant testing involves a few basic steps. Part to be tested is first thoroughly cleaned and then a liquid penetrant is applied to the surface. It is left for some time to sit on the surface while the penetrant penetrates the surface discontinuity, if there is any. After that, the penetrant remained on the surface is removed with water, cloth or thinner. Such removal cleans the surface of objects, but allows the penetrant to remain in the discontinuity. Then a developer is applied that works as a blotter and draws penetrant on the surface of objects and creates indication. Subject is inspected in order to detect indications that are visible due to contrast dye between a drawn penetrant and a background surface. After testing the surface is cleaned of residual penetrant and the protective layer is returned.
The advantages of this method are: low cost, portable equipment, it has a high degree of sensitivity; it gives instant results and requires minimum skills for performance and interpretation.
The disadvantages are that it requires a high degree of purity, it can detect only those defects that have a connection with the surface and there are no permanent written test results.
Ultrasonic testing
Ultrasonic testing method (ultrasonic inspection) is suitable for the examination of most metals, plastics and ceramics and defects on the surface or below the surface. Ultrasound examination requires that at least one part of surface near the surface to be tested is available. Examination of aircraft structures can be achieved by inducing ultrasonic waves on the object with the contact probe and receiving of reflected waves from that point. Reflection of ultrasonic waves is projected electronically into the tube of oscilloscope and is used to indicate defects.Ultrasonic testing method is a method of non-destructive testingthat uses high-frequency sound energy (above 20 kHz) for testing the structural integrity of the material. Ultrasonic waves can spread only in the medium. Precisely this fact is used to detect defects in the tested object. Ultrasonic waves at the medium border, as well as all other types of waves follow the laws of wave motion. For this reason, at the medium border, whether it is the wall of the tested object or irregularity, the reflection of ultrasonic waves and/or fracture, diffraction or other interaction between means and transmitted ultrasound energy will occur. Proper interpretation of ultrasound energy obtained from material testing can assess the condition of material and parameters of detected irregularities.
Radiographic testing
The radiographic testing of material, material irregularities are obtained in the way that the object of testing is aired with appropriate ionizing radiation. Radiographic testing will therefore show internal and external structural details of all portions of the material. It is a method that is used to check aircraft structure that is unavailable or unsuitable for the use of any other method
Recent research identifies interactions between corrosion and fatigue such that the presence of corrosion accelerates damage due to fatigue; thereby further reducing the total service life of an aircraft.
Fatigue occurs when a material is subjected to frequent loading and unloading. If the loads are above a certain threshold, microscopic cracks will begin to form, eventually a crack will reach a critical size, it will propagate suddenly, and the structure will fracture. The structure is tested to assess the situation and, if necessary, take corrective actions to prevent the occurrence of failures during operation.
Future recommendation
One of the ways of taking risk is:
1. Promote awareness through education of the safety measures to be taken during for enhanced security standards of the condition of the aircraft structures the main ones being the wings and fuselage.
2. Establish expectations through the publication of guidance material (Advisory Circulars and Staff Instructions).
3. Another very vital way of ensuring efficiency in the aviation sector is by regulating and enforcing the existing laws.
In a risk-based or risk-quantified approach to aircraft management, a distribution of crack sizes would be estimated, either analytically or based upon previous inspection experience, for a structure prior to an inspection. Key reliability problems faced in practice, such as:
1. Potential-cracking problems are revealed, and the aircraft is beyond its deterministic damage-tolerance limits.
2. Aircraft cracking has occurred to the extent that the deterministic-damage-tolerance derived inspection intervals need to be shortened in order to preserve safety.
3. Aircraft have been designed to be fail safe, but (widespread) fatigue damage has degraded the aircraft structure such that the fail safety of the structure has been compromised.

 

CONCLUSION
Significant quantities of energy are involved in transportation vehicles (aircrafts, trains, automobiles, etc.), as well as increasing engine operating temperatures, will enhance fuel efficiency. Obviously, new high strength, low density structural materials remain to be developed, as well as materials that have higher temperature capabilities, for use in engine components. Engineering materials will continue to play even more significant role in the current and future world. The factors that will influence this are found in economic/cost, environmental requirements, development trends, depletion of traditional materials, advances in research and market drives, etc. The importance of engineering materials in every aspect of life endeavor can, therefore, not be overemphasized. We ourselves are materials and so also is everything around us; to stop talking of and working with materials is to foreclose the essence of life existence. So a (Megerson:d.A.Mark, 2009)bright future is that of even more sophisticated, better and cost effective materials.

 

Leave a comment


Captcha

Related :-