Technical Report
For
Aero Structures – Properties & Performance
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Table of Contents
Table of Contents i
Revision History i
1. Repairing a cracked aircraft longeron 1 1.1 Abstract 1 1.2 Introduction 2 1.3 Damage Analysis of the Cracks in Longeron 2 1.4 Conventional Metallic Patch Repair 3 1.5 Bonded/fortified Patch Repair 4 1.6 Repair Design 4 1.7 Testing of the Repair Concept 6 1.8 Inspection of the bonded patch repair 8 1.9 Concluding Remarks 9 1.10 Repair Design Diagrams 10
2. Redesigning the Aircraft Wing 13 2.1 Introduction 13 2.2 How the Structure of an Aircraft being made 14 2.3 Design for Assembly and Manufacturing of Aircraft Wings 15 2.4 Types and Characteristics of Wings Assembly Method 15 2.5 Comparison of Wings Assembly Methods 17 2.6 Selection of Wing Assembly Method 18 2.7 Structural Wing Design & Analysis 18 2.8 Wing Structure (Spar) 19 2.9 Wing Structure (Ribs & Skin) 21 2.10 CFRP Composite Material for Constructing Wing Assembly 22
3. References 24 3.1 Books 24 3.2 Other Sources 24

Revision History Name | Date | Reason For Changes | Version | | | | | | | | |

Repairing a cracked aircraft longeron

A fighter was found with a fatigue crack on one of its longerons, which may eventually lead to a catastrophic failure. So, here in this report some of the techniques and possibilities are discussed as how to repair the cracks on the flange of the wings or longerons and what factors affect the repairing schemes and techniques.

Abstract

A two or three cracks or splits were detected in the flange of the upper left longeron of an aircraft. I first thought of how possible it is to apply the bonded patch repair scheme here instead of metallic patch repair. There are several other patching techniques for example bonded patch repair and the conventional reinforced patch repair. The repair needed to last 400 flight hours, and soon thereafter the particular aircraft would be reused for the operations. Titanium 6A14V sheet would be utilized for repair of the 2 mm thick 2024-T62 aluminum longeron flange. A symmetric bonded repair could be finished with a room temperature curing pressure sensitive adhesive. The attainability of the proposed repair geometry is dictated by performing range fatigue tests. After repair of the aircraft wing, occasional examinations would be done to check out on the de-bonding and crack propagations.

Introduction

Amid a stage review of the aircraft, splits or cracks are found out in the upper left longeron of the plane. The breakage occurred in the longeron flange for changing the access board or panel board of the left wing, at the area of the third clasp gap from the left of the board. There are two cracks on the aircraft longeron: one of the cracks is in longitudinal direction (outside the gap found) and the other one is transverse crack through the gap. The length of the longitudinal crack is around 70 mm furthermore; the length of the transverse crack is around 22 mm.
Because of the limited accessible space for repair a conventional or ordinary repair couldn't be implied to the cracked longeron. Removal and Replacement of the new longeron would require no less than 600 man hours. Further, conveyance time for another longeron could be dubious. Subsequently the possibility of a bonded patch repair could be considered. The repair needed to last at least 400 flight hours, since the particular aircraft could be resigned because of the limited life span of the aircraft.
This report gives an analysis of the cracks in the longeron, the fortified/bonded patch repair, performing range fatigue tests, and the patch review results amid the remaining existence of the aircraft.

Damage Analysis of the Cracks in Longeron

The flange is twisted to some degree in the upper left side of the longeron and the upper surface hints at showing the signs of friction. It is then noticed that the nut bolted to the flange lower side had a sideways orientation. These errors demonstrate a terrible fitting of the board on the longeron flange.
The particular access board may have been beforehand repaired with a moderately large stiffener that finished at the clasp hole where the cracks were discovered. It is no doubt that the access board with this large stiffener may have presented large static flexible stresses in the flange attributable to the bad fitting.

Operational fighter aircraft loads in combination with an assembly induction prompted flexible stresses most presumably is the reason of the longitudinal crack first. Accordingly, the transverse crack occurred from the latch opening attributable to dynamic malleable stresses in the longeron.
This conjectured spectrum of harm proliferation could not, however, be confirmed from the broken surfaces, subsequent to no piece could be cut off from the longeron if it is some how repaired.

Two Type of Longeron Repair
Conventional Metallic Patch Repair

As per the report of different aircraft repairing companies they suggests that a conventional metallic repair couldn't be implied on the cracked or broken longeron, for the most part due to the limited accessible space for repair.
Because of the severity of the breakage in a section that is a basic load path, we cannot use the technique of conventional metallic patch repair". The concerned longeron should be removed and substituted with the other one.
For removal and substitution of the longeron no less than 600 worker hours is almost required. Since the particular aircraft may retire from administration after around 400 flight hours, substitution of the longeron would be a costly solution. So therefore, different possibilities to empower a safe remaining operational life would be considered. The attainability of a bonded/fortified patch repair would be discussed.
So that is why installing a metallic patch repair is not the best repair scheme in that case because of the limitations.

Bonded/fortified Patch Repair

For a bonded/fortified patch repair to be doable, it would need to have the accompanying focal points: * Repair could be done at short notice, constraining the downtime for operational utilization of the concerned fighter aircraft. * The expenses would be little as compared to the expenses for removal and substitution of the longeron. * No extra holes would need to be penetrated. Such gaps cause a further debilitating of the structure which is a vital obstruction to a mechanical repair.
In determining the practicality of a bonded/fortified patch repair for the longeron which is cracked or affected somehow, the factors which should be considered are: * Design of a bonded patch repair that reestablish the quality of the broken auxiliary component. * Testing the design idea in research center tests representing the operational loads in the auxiliary component. * Process and methodology for inspection of the repair amid the operational life.

Repair Design

The composite patch repair focused on reclamation bonded/fortified patch does not remove the longitudinal breakage or crack in the longeron. So as to anticipate possible further development of the longitudinal crack, quit drilling at the tips of the longitudinal break is suggested.
Critical things/factors for the repair design are: repair patch material, repair patch measurements and type of adhesive. * Titanium 6Al4V toughened sheet with a thickness of 0.5 mm is chosen for the patch. The more slender the patch; the lower the miss match upon re-combining the access board. Moreover, Titanium 6Al4V has exceptional quality, consumption properties and also firmness. * For the dimensioning of the patch repair the following properties of material could be utilized: | σ.2 (MPa) | σ.ult (MPa) | E (GPa) | Longeron | 200 | 500 | 70 | Patch | 800 | 1000 | 110 |

The accompanying design rules are: 1. Stiffness × cross-segment of patch and flange must be similar. 2. Strength of the patch and flange must be about equivalent. 3. The adhesive bond section must be vast enough to exchange a definitive load in the flange made of aluminum. 4. A symmetric load presentation from the longeron towards the patch.

The proposed repair idea with a Titanium patch on both sides of the longeron flange must respond as discussed. For a 125 mm long and 25* mm wide patches this outcome in the accompanying analysis:

Firmness check E (Ai) × A (flange) = 70 × 2 × 22 = 3080 k(N) E (Ti) × A (Ti) = 110 × 25 × 2 × 0.5 = 2750 k(N) | Quality check Quality Al flange F (max) = 22 × 2 × 500 = 22 k(N) Quality T(i) patch F (max) = 25 × 2 × 0.5 × 1000 = 25 k(N) | Adhesive shear load at F (max) of Al flange Expect prominent adhesive length to be 45 mm Shear quality T adhesive = 22000 / 45 × 25 × 2 = 10 M(pa) |

Adhesive “Agomet” F 310 (Degussa) would be the prominent selection for bonding. This is “methyl methacrylate” based Gluing adhesive that cures at room temperature. The greatest shear strength is 30 M (Pa). The adhesive is most certainly not so important for the surface pre-treatment and furthermore, it has great gaps/holes filling properties. Accounting for knockdown elements owing to temperature and maturing, an adhesive shear force of 10 M (Pa) (from the shear stress examination) is considered worthy.

Testing of the Repair Concept

Keeping in mind the end goal to evaluate the performance of the patch repair, restricted fatigue tests should be performed on indented 25 mm wide strips of aluminium with and without a patch repair. Two specimens of open hole and two specimens of patch repair are prepared. A latch of 6 mm in diameter is connected in the focal hole of the repaired coupon specimen. The latch is torqued at 3 Nm, recreating the board connection to the longeron flange. In any case, load exchange happens solely through the bonded/fortified Titanium patches.

For examining fatigue and carrying out testing on it a forward fuselage bonding moment spectrum is ought to be used. However, directly usable sequence for the fatigue testing machine is rarely available. Measurement on aircraft demonstrates a pretty much comparable state of the forward fuselage and the wing root flexibility spectrum. Along these lines the LW-VAL load sequence is selected, this being illustrative for the wing root bending moment for a serious aircraft usage. For simplicity of testing the negative load levels in the succession are set to zero. To acquire a sensible testing time for the specimens with the focal hole, the fatigue tests are executed at 130 % of the LW-VAL load spectrum. This implies that the maximum gross stress in the specimens is 268.5 M (Pa).

The test program comprised of:

Open hole specimen | Specimen 1Specimen 2 | Static open gap strain test Open hole sequence fatigue test unsuccessful | Repair specimen | Specimen 3Specimen 4 | Fatigue testing for 15000 flights + lingering strengthFatigue testing for 3000 flights + lingering strength |

Prior to the lingering strength tests the patch repair specimens are reviewed for de-bonding with the Fokker Bond analyzer.
The static failure heap of the undamaged open gap specimen is 19.9 k (N), specimen 1. Fatigue testing at 130 % LW-VAL brought about an existence of 14300 flights, specimen 2. Specimen 3 demonstrates that the repair may survive 15000 flights after which a lingering strength of 23.1 k (N) should be acquired. NDI before lingering strength testing demonstrates local delamination at all the edges of the titanium patches. Specimen 4 may not hint at de-bonding after testing for 3000 flights. The remaining strength would be 25.3 k (N). Both specimens indicate malleable failure of the Titanium patches despite local fatigue prompted delamination for specimen 3.
Albeit no fatigue tests could be performed on bonded/fortified repairs after ageing, the repair idea appeared to be possible for an extra life of fighter aircrafts.

Inspection of the bonded patch repair

Before the Titanium patches are implied the disfigured attachment flange should be straightened to ensure great contact between flange and patch. The surface pre-treatment that comprise of Alumina Tri hydrate of the sanding of aluminum flange and Titanium patches. In the wake of bonding, the repair area should be furnished with the aircraft primer. An extra access panel should be settled to the structure utilizing fluid shimming to adjust for the patch thickness.
An inspection or examination strategy for the bonded/fortified repair should be set up for the remaining operational existence of the fighter aircraft. The inspection focuses on: * Visual inspection for paint breakage at the edges of the patch (showing de-bonding initialization). * Fokker Bond tester inspection for de-bonding of the patches. * Eddy Current inspection for cracks/breakage propagation from the penetrated off hole.
For keeping an eye on de-bonding the Fokker Bond tester Model 70 should be utilized alongside transducer 1414 and coupling liquid. The patch surface should be partitioned into areas of around 1 cm2. The conceivable Fokker Bond tester responses should be utilized as calibration signals amid the periodic inspections.
For crack detection and recognition the Nortec-19eII Eddy Current scope should be utilized with a Nortec PR/1 kHz – 100 kHz/a pencil test. The test frequency could be around 15 kHz. An alignment standard should be made to screen possible cracks/breakage propagation from the stop drilled holes.
Inspections after the bonded/fortified repair should be performed after 25, 90, 190, 300 and 400 flight hours. Indications of paint cracking at the edges of the patches, de-bonding or breakage propagation shouldn’t be found.

Concluding Remarks

In the present examination cracks or breakage in the upper left longeron of the aircraft are investigated. The fundamental cause for cracking is the existence of assembly stresses because of a despicably repaired access board. Since a conventional metallic patch repair couldn't be done, the exhortation is removal and substitution of the longeron. In any case, on the grounds that the aircraft would be resigned from administration after 400 flights hours, a financially savvy bonded/fortified patch repair is assessed and applied to the cracked longeron. Investigations amid the remaining operational life wouldn’t demonstrate any damage proliferation in the repair. In this way the present inspection demonstrates that adhesively bonded/fortified patch repairs can be extremely helpful and cost effective, particularly when the repair must be maintained just amid a limited operational life of the aircraft.

Repair Design Diagrams

* Repair Concept/Idea for the Creaked Longeron

Figure 1.10.1 Cross-Section of the Longeron

Figure 1.10.2 Drilled Holes in the Crack Patches

Figure 1.10.3 Effective Bonded/Fortified Region

Fatigue specimen for flight recreation Fatigue testing

Figure 1.10.4 Flight Simulation Fatigue Testing

Redesigning the Aircraft Wing
Aircrafts are built to meet certain predefined necessities. These necessities must be chosen so they can be incorporated into one airplane. It is impractical for one airplane to have all attributes; generally as it isn't feasible for an airplane to have the solace of a passenger transport and the maneuverability of a fighter plane. The sort and class of the airplane decide how solid it must be developed. A fighter aircraft must be quick, maneuverable, and prepared for assault and defense. To meet these necessities, the air craft is exceedingly powered and, has an extremely solid structure.
The airframe of an aircraft comprises of the taking after five noteworthy units:
1. Flight controls surfaces
2. Fuselage
3. Landing gear
4. Wings
5. Stabilizers

Introduction

This Report brings together the majority of the information contained in earlier sections and shows how it was implied on the configuration and design of the suggested assembly strategy for an aircraft wing subassembly. We will find in this section the use of Key Characteristics, the Datum Flow Chain, examination of limitations and resistances, and financial inspections, the assembly line of aircraft wing, the redesign process, life cycle for the re-designing of the aircraft wing. The process proposed in this report has not been implied on to the assembly described here, yet a considerable lot of the basic standards and qualities have been tried out on different substances.

How the Structure of an Aircraft being made

Structural or any specific component and configuration of an aircraft design is a subset of basic structural design as in general, including ships, land vehicles, scaffolds, buildings, and towers. All structures must be planned with consideration since human life regularly relies upon their execution. Structures are liable to limited and swaying stresses, the latter offering ascend to fatigues. Metal structures are liable to corrosion, and in the vicinity of stress and load the corrosion may accelerate.
Airplane structures are planned with specific regard for weight, for evident reasons. In the event that we could see underneath the interior fittings of an aircraft, we would see various lightening holes and gaps in the frame and also areas where the skins have been diminished by some sort of chemical milling. On some aircrafts, it is not find out the holes and gaps which are just the size of your hand with the thickness of their own, and many individual thicknesses are frequently found on a solitary skin. These regions vary in thickness by as meager as a millimeter or two, showing that impressive exertion is consumed to discover areas that are too daintily stressed. Such areas are intentionally made more slender to remove metal that is not doing its share of load bearing.
The principal aircraft had two wings made of light weight wood outlines with material skins, held separated by wires and struts. The upper wing and the struts gave pressure support while the lower wing and the wires bolstered strain loads.
In the 1920s, metal started to be utilized for air crafts. A metal wing is a case structure with the skins involving the top and base, with front and back framed by I-beams known as spars, interior fore-aft stiffeners known as ribs, furthermore, stringers also known as in-out stiffeners. In the balanced flight the lower skin is always in a strain condition while the upper skin is in pressure or the stage called as compression. Thus, this structure and configuration is alluded to as stressed skin construction. Amid turbulence, upper and lower skins can encounter both compression and tension. This case structure can bolster the aforementioned moments, making single wing aircrafts conceivable. The end of the struts and wires so drastically decreased air drag that air crafts could fly twice as quickly as before with the same engine.
Typical or conventional Metal Skin Aircraft Fuselage Assembly structure comprises of a skin to which have been bolted longitudinal stiffeners (along the 34 ft. bearing) called stringers. Alongside the circumferential of 22 Ft. bearing, there are stiffeners called frames. Every stringer-design crossing point is joined by a little piece called a clasp.

Design for Assembly and Manufacturing of Aircraft Wings

* Sophisticated items include an extensive number of singular components and subassemblies. * 70 to 80 percent of the expense of assembling an item is resolved amid the design stage * A sane design for simple and minimal cost assembly is the choice of the most suitable technique for assembling these items. * So, a configuration Engineer ought to be concerned with the simplicity and expense of assembly. * Thus, the idea of configuration for assembly (DFA) risen.

Types and Characteristics of Wings Assembly Method

Manual Assembly: * While assembling manually, the operations are completed manually with or without the guide of straightforward, general purpose apparatuses like screwdrivers and forceps/pliers. * Individual parts are given to the workbench either physically or by utilizing mechanical gear for example, parts bolsters or exchange lines after that they are manually assembled. * This assembly technique is described by its versatility and adaptability. * Per item cost of assembling is almost the same. * Independent of the generation volume.

Automatic Assembly: * Fundamentally alluded to as altered automation or the Detroit sort. * Either programmed feeders and synchronous indexing machines or non-synchronous machines where parts are taken care of by free-exchange devices are utilized. * Machines are committed for the generation/assembly of the items. * These frameworks do not have any adaptability to suit tangible changes in the designing and configuration of the product. * Requires a huge scale capital investment, and in addition significant time and designing work before actual production can be begun.

Programmed Assembly Using Robots/Robotic Assembly: * The amount of manufacturing amount is greater than that of a manual assembly framework however lower than that of a programmed assembly framework (Settled automation).

Basic types of Robotic Assembly:
1. one-arm robot operating at a solitary workstation that incorporates parts feeders, magazines, and so forth.

2. Two robotic arms operating at a solitary workstation * A programmable controller (PLC) is utilized to at a time synchronize and control the movements of the two arms. * Referred as an automated assembly cell and just like FMS cell.
3. Multi-station robotic assembly framework * Capable of performing a few assembly operations at a same time. * Can perform diverse assembly operations at every station. * High adaptability and flexibility to design changes.

Comparison of Wings Assembly Methods

* Manual assembly requires the minimum capital investment followed by the two least complex types of robotic assembly. * Automatic framework compares to different stations automated assembly framework with unique purpose frameworks and machines requires more capital investment for a vast production volume however for the moderate manufacturing volume it requires little capital investment. * Assembly cost per product is steady for manual assembly. * Assembly cost per product diminishes directly with expanding production volume for programmed assembly utilizing special purpose machines. * For the situation of automated assembly, the assembly cost per item diminishes with expanding production volume, however turns out to be less efficient after surpassing the yearly production volume at certain point.

Selection of Wing Assembly Method

Factors influencing selection of an assembly technique: * Annual generation volume (or production rate) * Cost of assembly * Number of solitary components to be amassed in a product * Number of various renditions of a item/s * Availability of labor (with expense consideration) * Payback period * These elements are interactive * Impossible to have a solitary mathematical relationship between these factors.

Structural Wing Design & Analysis

The structure of the wings is a region where huge weight can be taken off by picking the right material, yet the assembling strategies additionally must be weighed into the material choice. The primary material considered is a froth/foam which would have had carbon fiber spar and a carbon fiber skin. This material has the lower weight however it is additionally an extremely muddled assembling process. The second material that is considered is balsa wood for the ribs, pre-assembled carbon fiber tube for the spar, “Mono Kote” which is a plastic sheeting for the skin. This is a lightweight mechanism; however the assembling for it will be much less complex. The choice to utilize the second set of materials as a result of simplicity in assembling should be taken more into consideration.
Flight Envelope
The most extreme load on the airframe is thought to be in the hypothetical instance of the aircraft going at its greatest velocity, and after that suddenly pitching up to its stall angle. This greatest instance of around 2 times the basic load amid enduring level flight compares to a dispersed load of around 46 pounds on the front wing. The flight envelope is determined by finding the normal aerodynamic stress at speeds going up to the top rate of the aircraft if the control surface is to be completely avoided. This flight envelope is utilized when performing all wing structural design and analysis.

Wing Structure (Spar)

At first the spar design starts with an estimation of the size and loading in conjunction with aggregating a list of desired conduct attributes. An estimate for the loading and size originates from the requirements set by the SAE competition rules for the normal class competition. Those being, that the most extreme weight of the loaded plane, including fuel and payload, should not surpass 55 lbs. From this a basic power balance for the plane amid in high velocity flight conditions implies that the load on the wings should be 55 lbs. This is a rough evaluation utilizes for starting calculations; more exact design would occur later utilizing the most extreme gross takeoff weight multiplied by the load factor dictated by the flight envelope.

The maximum wingspan accessible is estimated utilizing the dimensional restrictions suggested by the competition guide lines and rules. The guidelines now require that every plane should not surpass a 175 inch sum of the stature/height, length and width of the plane, barring the prop-length. From those reading the main stature is calculated dependent on the clearance for already existing propeller length, giving an inexact tallness of 18 inches. This lefts a maximum wingspan of 9 feet, and gives the maximum starting length for the spar design limitations. The accompanying list comprises of the objectives that are resolved to have ideal execution of the spar: * Maximize strength * Maximize quality * Minimize weight * Maximize rigidity * Manufacturing simplicity

Maximizing quality and strength is important to withstand the most lift power so as to lift the most weight, which is the whole purpose of the competition. This strength and quality constraint is mainly tended to by material choice however certain configurations additionally influenced this objective. Due to the long wingspan, in order to diminishing any vibrations and fluttering of the wings, the spar design should be as unbending or rigid as possible.

Other than the general configuration component of security of 1.2 expanded the conservative nature of the estimations by concluding that model of the spar should base on the presumption that all the lift would be consumed by a individual spar despite the fact that an secondary spar will be utilized to handle the torque experienced by the wing and servo/control surface arrangement alongside expanding the general rigidity nature of the structure.

The expansion of a bracket that would retain all the flexibility of the spar diminishes the effective bending length of the beam to 46 inches, which would be considered in both of the final analysis estimation methods. It ought to be noticed that the analysis results are to some degree overstated in light of the fact that the addition of the optional spar and the ribs with the MonoKote covering will build the rigidity nature of the wing and decrease the redirection.

Bracket:
The design of the fundamental spar bracket depends on the accompanying objectives: * Fit for retaining the bending stress from the principle spar * Does not transmit spar stresses to the fuselage association * Effortlessly fabricated

Wing Structure (Ribs & Skin)

The ribs will be developed out of bits of balsa wood slice to the picked airfoil shape. The shapes would be cut out by following the shape off of a bit of paper utilizing a pen and after that could be cut out of the primary sheet of wood utilizing a basic extremely sharp razor blade. After the wing is assembled it would be secured with MonoKote by just attaching the sheet down and after that applying warmth to it with an iron. The MonoKote gives a few favorable circumstances over the composite layup. To start with, the MonoKote has an easier developing process then the composite layup. Also, the composite material is difficult to repair if there is an issue with it, yet with the MonoKote every one of that must be done is reapplying the warmth. In the event that there are wrinkles it is simply apply heat, for a hole or a gap there is pressure initiated MonoKote to apply in the field, and after that a patch can be made of the main MonoKote which is then applied with warmth and heat at certain degrees.

This material selection is made in view of the simplicity of manufacturing and assembling, and it takes into account differing the dispersing between the ribs as it looks fit. This dispersing and spacing between the ribs can be dictated by a few strategies including yet not restricted to the stress that the ribs will encounter and the pressure required from the skin to hold it taught. After literature search it is resolved that underlying failure because of extensive rib spacing and gaps would be clasping of the MonoKote skin. This failure will happen before the spacing turns out to be sufficiently huge that the ribs will separately bolster enough force to bring about failure. A dividing of 3" should be chosen, as models of comparative size have somewhere around 2 and 4 inch spacing.

CFRP Composite Material for Constructing Wing Assembly

CFRP Composite materials are lightweight, solid materials utilized as a part of the assembling of various items utilized as a part of our everyday life and it is a basic component as well in constructing the wing assembly of the aircraft. Carbon Fiber Reinforced Polymer Composites is a term used to portray a fiber fortified composite material that uses carbon fiber as the essential structural element. It ought to be noticed that the ‘P’ in CFRP can be used for the plastic instead of polymer.
By and large, CFRP composites use thermosetting pitches, for example, epoxy, vinyl ester, or polyester. Albeit thermoplastic pitches are utilized as a part of CFRP Composites, "CFRTC" frequently pass by their own particular acronym, CFRTP composites.
At the point when working with composites or inside of the composites industry, it is critical comprehend the terms and acronyms. All the more critically, it is important to comprehend the properties of FRP composites and capacities of the different fortifications, for example, carbon fiber.

Advantages and Disadvantages of utilization of CFRP over Aluminum Alloys:
Advantages:
Light Weight - A customary fiberglass reinforced composite utilizing consistent glass fiber with a fiber of 70% glass (weight of the glass/total weight), will usually have a thickness of .065 pounds for every cubic inch.
In the interim, a CFRP composite, with the percentage same as 70% fiber weight may regularly have a thickness of .055 pounds for each cubic inch.
More Stronger- Not just are carbon fiber composites lighter weight, however CFRP composites are much more grounded and stiffer per unit of weight. This is genuinely true when contrasting carbon fiber composites with glass fiber, yet significantly all the more so when contrasted with metals.
At the point when comparing CFRP composites with aluminum alloys, one of the lightest metals utilized, a standard presumption is that an aluminum structure of equivalent quality and strength, its weight would be around 1.5 times greater than the structure of the carbon fiber.

Disadvantages:
Cost/Value - Although astounding materials, there is a motivation behind why carbon fiber is not utilized as a part of each and every application. Right now, CFRP composites are cost restrictive in numerous cases. Contingent upon the present economic situations (supply and demand), the kind of carbon fiber (commercial grade vs aerospace), and the cost of carbon fiber can fluctuate significantly.
Crude carbon fiber on a cost per pound, can be anyplace between 5-times to 25-times more costly than fiberglass
Conductivity - Carbon fiber is greatly conductive, while aluminum steel alloys are very less conductive. Numerous applications use glass fiber, and can't utilize carbon fiber or metal, entirely in view of the conductivity.
In spite of the fact that the expense of CFRP composites still stays high, new mechanical progressions in assembling are keeping on considering more savvy items. Ideally, in our lifetime we will have the capacity to see financially savvy carbon fiber utilized as a part of an extensive variety of consumer, modern, and automotive applications.

References
Books
Mechanics of Aero-structures 1st Edition by – Sudhakar Nair
Analysis and design of flying vehicles by - E.F.Bruhn
Modern Aircraft Design, Volume 1, 6th Edition by – Martin Hollmann
Aircraft Design: A Conceptual Approach, Fourth Edition - by D. Raymer
Other Sources http://adg.stanford.edu/aa241/AircraftDesign.html http://www.allstar.fiu.edu/aero/flight12.htm http://www.f-16.net/f-16-news-article4913.html https://www.flightglobal.com/news/articles/pictures-manufacturing-defects-caused-cracks-that-downed-usaf-f-15-220799/
http://www.aeroncapilots.com/98CE121AD.htm