Design and Certification of an Aircraft Major Modification:

Ka-Band Satcom System for Airbus A320 Family

Bernardo Manuel Machado Duartebernardo.duarte@tecnico.ulisboa.pt

Instituto Superior Tecnico, Universidade de Lisboa, Portugal

November 2019

Abstract

With the increase of the number of aircraft in the sky in the last decades, a new business productemerged, the aircraft modifications. With the support of Jet Aviation, a company specialized in thefield, this work goes through all the necessary steps to apply for a Supplemental Type Certificate(STC) to perform an installation of a ka-band satcom system, classified as a major modification, inthe primary structure of an aircraft, by complying with the EASA and FAA regulations. The mainsteps covered are the design development, certification requirements, analyses, installation, experimentaltests and required documentation. The structural provisions for this system are developed in-houseand are based on an ARINC industry standard. The gains of using this standard in the design areapproached. The main object of this thesis is the development of a parametric optimization study tothe structural provisions in terms of weight reduction and fatigue life. To conduct that, three structuralmodifications in three different components of the initial design are defined and seven new designsderived from the Jet Aviation one are created. The seven hypotheses are subjected to a static analysisperformed by a combination of FEM simulation and analytical methods, to verify their structuralintegrity. Furthermore, a fatigue analysis is conducted to the critical parts to estimate the new fatiguelife of the designs. As result, is determined the optimal design between the seven in study.Keywords: Aircraft modifications, Ka-band satcom system, Certification, Static Analysis, FatigueAnalysis

1. Introduction

With the fast growing number of aircraft flyingsince the introduction of jet aircraft for commercialuse in the 1950s [7], new opportunities of businessproducts emerged, including the aircraft modifica-tions. An aircraft modification is defined by beinga change/alteration or a repair. A change is the ac-tion of adding new features to the aircraft, where as,a repair is the action of re-establishing the originalstrength of a damage area. Furthermore, a changecan be classified as major or minor, depending on ifit might appreciably affect weight, balance, struc-tural strength of the aircraft [8].

Any modification performed to an aircraft mustbe certified by an aviation authority. The main or-ganizations responsible for this task are the Euro-pean Aviation Safety Agency (EASA) in Europeand the Federal Aviation Authority (FAA). Thereare different ways to certify a change. One of theoptions goes through the creation of a SupplementalType Certificate (STC). This option is more expen-sive compared to the other options and normally isused when there is an interest in performing the

same modification several times.

The elaboration of this study was developed dur-ing my traineeship in a company named Jet Avia-tion, which provides me all the support to conductit. Jet Aviation is a company of the business avia-tion industry sector which provides several differentservices, including engineering services. From theseservices, the one addressed in this work is the in-stallation development of ka-band satcom system[10].

The ka-band satcom system is the state of the artin aircraft communication via satellite. This systemallow to have high speed data on board in a widearea of the globe. It is composed by the systemswhich provide the service and the structural partwhich is the mechanical interface between the an-tenna and aircraft. The installation of a ka-bandsystem is classified as a major change due to thefact that it will affect the structural strength of theaircraft [11].

Jet Aviation started the development of this sys-tem for the Airbus A320 family to afterwards cre-ate an STC, where are included the A319 and A320

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series aircraft. The development of the structuralprovisions of this system is the main base for thework done in this article.

2. ARINC 791 Standard

The structural provisions of ka-band system devel-oped by Jet Aviation have as a background the in-dustrial standard known as ARINC 791.

The ARINC characteristic 791 is one of severalstandards published by SAE-ITC. This organiza-tion promote the ARINC industry activities whichpromote the safety, efficiency and cost effectivenessin aircraft operations [5].

This characteristic provides the full definition forthe ku and ka-band satellite communication sys-tem in terms of physical installation, aircraft inter-faces and electrical interfaces and functional equip-ment as well. The main systems which composethe ka-band system given in the standard are theAntenna Aperture (AA) which receives and trans-mits the radio frequency signals, the Ka-band Ra-dio Frequency Unit (KRFU) which converts the sig-nal from the AA, the Ka-band Aircraft NetworkingData Unit (KANDU), which controls the KRFU,the MODMAN which is composed by the modemand the manager, and the Aircraft Personality Mod-ule (APM), where all the configurations for the sys-tem in each aircraft are saved.

Regarding the aircraft interface, this standard de-fines that the antenna must be attached to the fuse-lage by seven fittings fixed on its skin. The positionof the seven fittings is also established by it andthis provides the initial point to start the design ofthe internal provisions, which include the fittingsas well. Furthermore, the ultimate loads for thefittings design is given by the standard.

With the employment of the ARINC standard791 in the development of the structural provisionsof the ka-band system there is a huge earned value.Firstly, the standard is valid for all types of aircraftand independent of the manufacturer [5], this meansthat, the developed solution can be adapted to anyairframe in a short time. Secondly, the equipmentsrequired for the installation are not exclusive, differ-ent manufacturers equipments can be fitted in thestandardized provisions [5], which signify the end ofthe dependency on only one supplier. And thirdly,the fully definition of the mechanical interface be-tween antenna and aircraft provide a way to startdesigning not from the scratch.

3. Ka-band System Structural Base Design3.1. Design Constraints

Before detail design stage starts, it is necessary tobe aware of all the conditions which will constrainthe design.

First of all, the certification requirements. Ev-ery aircraft has a certification basis linked to itself,

which shows that aircraft complies with a specificamendment of the certification specifications (CS).For the modification in study, Jet Aviation electedto comply with CS25 amendoment 22, which is thecertification basis of the target Airbus A319. Thecertification paragraphs which must be shown com-pliance by the structural provisions are presented intable 1. The description of each paragraph is foundin [2].

Title ParagraphLoads 25.301Factor of Safety 25.303Strength and Deformation 25.305Proof of Structure 25.307Flight Loads - General 25.321PressurizedCompartment Loads

25.365

Emergency LandingConditions - General

25.561

Material StrengthProperties and Design Values

25.613

Fitting Factors 25.625

Table 1: Certification paragraphs applicable to thestructural modification [3].

Secondly, the location where antenna will be in-stalled is analysed. The optimal position is found ontop of the fuselage in the upper crown area alignedwith the vertical plane of aircraft’s symmetry, dueto the reduction of anti-symmetrical flow pertur-bation. In terms of the longitudinal location, sev-eral factors play an important role. These are thecritical aerodynamic sections, the existing externalequipment, the inside aircraft’s environment andthe non obstruction light condition. Regarding thefirst factor, the area between the nose and end ofthe wings is considered aerodynamically critical dueto any flow perturbation in the region can have aconsiderable impact on the conditions of the airflowincoming on the wings. Thus, the installation loca-tion chosen for the current aircraft was on the rearpart of the fuselage, after the wings.

3.2. External Supplied ComponentsSome structural components are not designed byJet Aviation, these are bought from suppliers. Theradome, adapter plate and fairing are the threeparts which are bought externally. The radome isthe component which covers the outside antennaequipment and protect it from the external agents,such as, dirt, hail stones, water, de-icing fluid andbirds [5]. The adapter plate is the antenna base andalso the mechanical interface between the radome,fairing and the Jet Aviation’s designed structure.These two components fit in all aircraft types.Lastly, the fairing is the part which is specific for

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each aircraft type and its mainly function is to pro-vide the drainage required to avoid trap water andavoid corrosion [5]. Figure 1 presents an explodedview of these three components plus the AA andKRFU.

Figure 1: Exploded view of external equipment andoutside antenna equipment [1].

3.3. Jet Aviation’s Structural Provisions DesignTo start the detail design of the structural pro-visions, firstly, the environment of the A319 wasbuilt with the main structural components of theairframe, such as, frames, stringers, skin and shearclips. Concluding the environment, the provisionswere designed on top of that. The developed solu-tion is made of seven external fittings which securean adapter plate that supports the antenna mount-ing plate and radome. Each fitting is attached tothe aircraft external fuselage surface through a dou-bler. A mounting gusset is installed inside the fuse-lage along with each external fitting, supported byinternal intercostals within the fuselage supportedby the fuselage frames. Different type of fastenersperform the connection between this components.Figure 2 illustrates the final solution of the A319,reusable also for the A320.

Figure 2: Jet Aviation structural provisions 3Dmodel installed in the upper lobe fuselage with iden-tification numbering of each fitting.

To sum up, the structural components whichcompose the design are presented in the followinglist.

• External fittings

• Doublers

• Mounting gussets

• Intercostals

All these parts are made of aluminium alloys withdifferent properties and some of them are machinedand others are sheet metal bended. In order to per-form the sizing of these parts several failure modeswere identified to be analysed and checked for struc-tural integrity.

4. Parametric OptimizationThe parametric optimization began with the cre-ation of seven candidates for the optimal design.Followed by the selection of the critical load casebetween thirteen initial load cases. Finally, themethodologies used for the analyses are introduced.

4.1. HypothesesTo create seven new designs, three modifications inthree different components were elaborated in or-der to combine them and obtain the seven differentcombinations for the new designs. The fittings, dou-blers and intercostals were the components whichwere found more interesting to modify. These mod-ifications are explained in the follow paragraphs.

• Fittings (F) - In order to reduce weight, theoverall thickness of the fittings was reducedfrom 0.346 to 0.189 inches, except the areaaround the lug which remain 0.346 inches. Fur-thermore, the thickness of the fittings’ base wasdecreased to 0.197 inches. Figure 3 illustratesthe modifications done in the fittings.

Figure 3: Illustration of fittings modification - fit-ting one.

• Doublers (D) - The usual process for doublerdesign is driven the structural repair manual ofthe aircraft which states that the doubler thick-ness must be one gage (0.008 inches) thickerthan the skin panel in touch. Here, will bestudied the case where the doubler’s thicknesswill be the same of the skin panel. Thus, themodified doubler will be 0.063 inches thick andnot the initial value of 0.071 inches.

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• Intercostals (I) - In Jet Aviation design for theA320 family, the intercostals were sized with athickness of 0.1 inches, where as, for the Boeing737 family aircraft, these have only 0.08 inches.These two families of aircraft are in a similarweight category, so the intercostals modifica-tion proposed in this study is the reduction oftheir thickness to the 0.08 inches.

Therefore on table 2, the seven hypothesis aredefined by the following combinations of modifica-tions.

Design Hypothesis

1 F

2 D

3 I

4 F+D

5 F+I

6 D+I

7 F+D+I

Table 2: Definition of the seven hypotheses for thestudy.

4.2. Load CaseIn order to perform the static and fatigue analyses,the load cases for each analysis must be defined.There are thirteen load cases which were used inthe substantiation of the initial design. This set ofload cases are composed by gust, fatigue or crashload cases, or rapid decompression of the radomeand fuselage burst pressure. All situations whichcan happen during a flight and thus, the ka-banddesigned provisions must withstand them. Theseload cases are a combination of the inertia of thedesigned structural provisions plus the equipments,the bending and shearing of the fuselage barrel inthe target section, and the aerodynamic and pres-surization loads. The cabin operating pressure, andstatic flight loads envelope are extracted from [6].

For the static analysis the load case selected forthe study was the radome rapid decompression load,which can be assumed as a good approximation ofthe most critical load case. For the fatigue analysisthe fatigue load was used. Table 3 presents thecomposition of each load case.

4.3. MethodologiesRegarding the static analysis, this was performedin two steps. Firstly, the reaction loads on the fit-tings’ lug were extracted from the results given bythe FEM model developed in FEMAP for all thehypotheses. Secondly, with the extracted loads, an-alytical methods were employed to determinate thesafety margins for all the parts of the design. For

Inertia

Bending &

Shear

Fuselage

Cabin Pressure Radome

1.0g

Down

1.0g

Down

DP = 8.99 psi

Operating

Pressure

3 Psi Radome

Over Pressure

+ CFD Cruise

Load Case

1.3g

Down

1.3g

Down

DP = 8.99 psi

Operating

Pressure

CFD Cruise

Load Case

Table 3: Load cases composition for the radomerapid decompression and fatigue load extractedfrom [3].

this study, the consider failures loads used for thesafety margin calculations are only presented for themodified parts.

Therefore, the failure modes consider in the anal-ysis for the fittings, doublers and intercostals werethe following:

Fittings

• Lug analysis

• Fitting strength

• Base bearing

• Von mises stress

Intercostals

• Bearing

• Outer/inner flangesstress

• Web shear stress

• Von mises

Doublers

• Bearing

• Von mises stress

The global methodology used to determine thesafety margin for each of the failure modes was sim-ilar for all. Equation 1 was the core equation usedin all the calculations of the safety margin.

MS =F

FF × f− 1 (1)

MS is the safety margin, FF the fitting factorimposed by the certification paragraph 25.625, Fis the allowable stress of the material for a specificdirection (L or LT) under the ultimate or yield con-dition and f is the maximum stress found in eachanalytical analysis. For the different failure modes,the computation of this value is performed by differ-ent sub-methodologies which are presented in detailin the complete document.

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Regarding the fatigue analysis, this is based onthe material Wohler curves, also known as stress-life curves (S-N), which relates the maximum stresswith the fatigue life, in cycles. For this analysis isassumed that one cycle is equivalent to one flightby considering a constant amplitude Ground-Air-Ground (GAG) loading cycle [9]. The goal of thisanalysis is justify that the fatigue life is not a limit-ing factor for the threshold inspection. Due to thehighest load condition, the doublers are the criticalcomponents which must be subjected to the study.Figure 4 illustrates the S-N curves for the doublersmaterial - Aluminium alloy 2024-T3. In addiction,also the fatigue life of the modified fittings is studiedin order to compare with fatigue life of the initialfittings.

Figure 4: S-N curves for aluminium alloy 2024-T3from [4].

Therefore, to extract the fatigue life from the cor-relations of the graphic in figure 4, firstly is nec-essary to compute the stress concentration factorwhich is estimated by the theory of severity fac-tor (SF) times a discrepancy factor of 1.2 for thedoublers. For the fittings case, this factor is com-puted by the quotient between the maximum prin-cipal and nominal stress extracted from the FEManalysis. Secondly, the local stress concentrationfactor and net section stress are computed to fi-nally determinate the equivalent stress which is therequired parameter in the correlation to obtain thefatigue life. For the fittings the equivalent stress issimply the maximum principle stress extracted fromthe FEM analysis. Equation 2 presents the correla-tion to obtain the fatigue life of the doublers.

Log(Nf ) = 8.3 − 3.3log(Seq − 8.5) (2)

Finally, to obtain the required value is necessaryto divide the Nf by a scatter factor of 5 used by theOEMs. This value is the fatigue life number cycles

used to compare with the initial values. The detaildescription of the methodology is presented on thecomplete document.

5. Results

Several results were obtained from this study.Firstly, it is presented the obtained results in termsof weight of the several hypotheses. Secondly, theresults of the static analyses are shown. Thirdly,the fatigue life cycles of the modified componentsare presented as part of the fatigue analysis results.Finally, a combination of all the results are per-formed to find out the optimal hypotheses.

5.1. Weight

Regarding the weight results, with the modificationperformed in the fittings, doublers and intercostalsto define the seven design hypotheses as a combi-nation of these three changes, the weight reductionflow with all the hypotheses are presented in table4 and figure 5.

Table 4: Values of weight reduction percentage forthe seven hypothesis.

Figure 5: Percentage of weight reduction of initialJet Aviation provisions.

The results presented in table 4 shows that hy-pothesis number seven, which includes the threemodified components, presents the highest weightreduction, a reduction of 10.4% of the total weightof the initial design. In addition, it also shows thatthe modified component with the larger impact onthe weight reduction is the intercostals, due to thelarge number of times that this component is usedand also, to its considerable initial weight comparedto the other two modified parts. Following the samelogic, accounting only for two modified components,the hypothesis with the greater result is the hypoth-esis number five, which combines the modified in-tercostals with the modified fittings, which are the

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second component with more impact on the weightreduction.

To sum up, the hypothesis number seven is thelightest design from all the hypotheses. The staticanalysis results will confirm if the structural config-uration of this hypothesis maintains the structuralintegrity by continuing to have positive safety mar-gins in all the design components.

5.2. Static

Before performing the analysis, the reactions loadsin the fittings in the three directions were extractedfrom the FEM analysis. Figure 6 presents the valuesobtained for the Z direction, which gives a goodrepresentation of what happens in the other twodirections.

Figure 6: Fittings reactions in Z direction.

These results show that for all the configura-tions the reactions in the fittings are almost similar.Comparing the reactions from the seven hypothesisagainst the reactions of the initial design, it showsthat the reactions suffer short alterations, at unitsscale, as it is possible to confirm with the nearlyhorizontal lines in the graphics of figure 6. Theseresults show that the geometry modifications per-formed in the selected parts had an insignificant im-pact on the global structural systems rigidity, whatexplains the shorts changes in the reactions.

From the second step, the safety margins werecalculated for the specific failure modes for the fit-tings, doublers and intercostals. Below are pre-sented only the obtained critical results for each ofthe three components.

• Fittings

Table 5: Safety margins results of fitting 3.

Observing the results from table 5, firstly, itshows that all the fittings are able to withstand theapplied loads on all the hypothesis. Secondly, it ispossible to confirm the safety margin reduction insome of the failures modes. One example is the fit-ting strength due to the bending, which the safetymargin goes from 2.56 in the initial design to 1.51 inthe last hypothesis on fitting three, thus this is thecritical value obtained and the fitting three is thecritical fitting, however, it is still in a safe zone. Thecritical values for the fitting three are highlighted inblue and bold. Lastly, looking at all the values ofthe seven hypotheses, the accentuated safety mar-gin reduction always occur in the hypotheses wherethere are the fittings modification.

• Doublers

Regarding the doublers, the analyses performedshowed that the doubler three is the most critical.Table 6 presents the safety margins results for thatdoubler.

Table 6: Safety margins results of critical doubler -Doubler 3.

Observing the results of table 6, the doubler threeis the component with the lowest safety margin,with a value of 0.02 for the hypothesis seven. Inspite of being very close to the ultimate allowableload, it is still under it, so it remains in the safezone. Furthermore, it is known that the load caseapplied is conservative, so it is concluded that thedoubler has sufficient strength.

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• Intercostals

Table 7: Safety margins results of intercostals infitting 7.

Observing the results in figure 7, firstly, it showsthe intercostals in fitting seven presented the lowestsafety margins, however the analytical method useddoes not account for the cut-out in intercostals offitting 3 and 4. Thus, the critical intercostals are infitting 3 and 4, which are shown by the FEM results.Secondly, it shows that all the intercostals are ableto withstand the applied loads on all the hypothe-sis. And lastly, it is possible to confirm the safetymargin reduction in some of the failures modes inthe successive hypothesis.

5.3. FatigueRegarding the fatigue analysis, the follow resultswere obtained for the critical fitting and doubler.

• Fittings

Table 8: Fatigue lifes number of cycles of fitting 3for the initial design and seven hypotheses.

Observing the results in figure 8, firstly, it showsthat the fitting modification decreases the numberof cycles in 93% due to the increase of the maxi-mum principal stress in the base of the fitting. Onthe other hand, the number of cycles for the mod-ified fitting is still acceptable, because is below tothe limit of 24000 flight cycles imposed as thresholdby Airbus for the aircraft models A319 and A320.This value is obtained by equation Nth = 0 : 5DSGprovided in the Airbus structure training manual,where DSG is the Design Service Goal and is equalto 48000 flight cycles for the mentioned aircraft.These information is given in the Airbus fatiguestress manual.

• Doublers

Table 9: Number of cycles of fatigue life for doubler2 for the initial design and seven hypothesis.

The results shows a reduction in the number ofcycles for all the hypothesis, having the maximumreduction in the hypothesis where the doubler withreduced thickness is employed. In addition, it alsoshows the impact of changing the intercostals andthe fittings in terms of doublers fatigue life, witha reduction of 10% and 6% in the number of cy-cles, respectively. In spite of that reduction, allthe hypotheses have the number of cycles above thethreshold imposed again by Airbus of 24000 flightcycles.

Cross checking the obtained results for the fit-tings and doublers, it shows that with the fittingmodification, this component became the criticalone in terms of fatigue life, due to the huge reduc-tion of the number of cycles. Therefore, the fatiguelife of hypotheses 1, 4, 5 and 7 is driven by the fit-tings and not the doublers as in the initial design.The other hypotheses continue to be driven by thedoubler and does not introduce a huge reductionin the number of cycles. Figure ?? summarize thecritical values of the number of cycles for all hy-potheses.

5.4. Optimal Hypothesis

Once all the results are presented, it is time to de-cide which is the optimum hypothesis to go for.This decision is a compromise between several re-quirements, such as, the certification, the businessand the company internal requirements. The busi-ness ones are traduced by the customer require-ments, where there will be two possible scenarios.Or the customer is an airliner or a private one. Incase of an airliner as a customer, it is known that thegoal of an airline is to make profits in their flights,so less weight introduced in the aircraft means lessfuel consumption, which reflects in less operationalcost, so for an airliner the main requirement is theweight reduction. On the other hand, for a privatecustomer which is not worried about weight but isconcerned about extra costs, the inspections inter-val are the main requirement, the design providedmust guarantee that the customer will not have ex-tra costs in terms of maintenance due to shorts in-tervals of inspection and the aircraft must be onground only because of the antenna. Regarding theinternal requirements, these will influence the de-veloped design. Examples of these requirements arethe capacity of company to produce a certain designand also if the company has the knowledge requiredfor the development of that.

To support the decision normally a graphicnamed pareto front is produced, where for two vari-ables are plot the several hypotheses to be easy tocross check the results and confirm which is opti-mum decision according to the weight given for each

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variable. Figure 7 presents the graphic with thecritical fatigue life number of cycles which driveseach hypotheses versus the weight reduction.

Figure 7: Pareto Front - Critical fatigue life’s num-ber of cycles VS Weight reduction.

Observing the graphic above, it is possible to seethat there are two groups of sets of hypothesis, oneclose to the 1% weight reduction and another closeto the 9%. The first group is composed by hy-potheses 1 (F), 2 (D) and 4 (F+D). Furthermore,hypothesis 1 and 4 in this group has a reduction inthe fatigue life comparable to other hypothesis withhigher weight reduction percentage. Thus, thesehypotheses are far to be the preferable ones. Thesecond group is composed by hypotheses 3 (I), 5(F+I), 6 (D+I) and 7 (F+D+I). From these hy-potheses, the number seven is the one which im-proves more in terms of weight reduction, howeverhas a reduction impact of 59.8% on the number ofcycles. The optimum option can be hypotheses 6which has 9.2% of weight reduction and the impactin the reduction of number of cycles is lower thanthe hypothesis seven, 18.8% for this hypothesis .

Regarding the last requirement, the certification,all these hypotheses of design are not certifiable yetbecause it is necessary to prove that the hypothesiscan withstand all the other mentioned load casesand for this study it was only considered one of theload cases which is a good representation of the crit-ical load cases but is not enough. In addiction, adamage tolerance analysis must be performed in or-der to find the inspection intervals, performing onlya fatigue analysis is not enough again for certifica-tion. In spite of that, this initial study gives goodindications on which improvements can be done inthe initial design in order to achieve higher qualityvalues.

6. Design Initial ReleaseReaching to the final stage, once the structural anal-yses are completed, the solution design is convertedinto 2D drawings. There are the installation draw-ing and drawings for each component of the design.When all these are completed and released by theengineering, this moment is known as the Initial

Release. The next step is the installation when theparts have been already manufactured by the pro-duction department. During the installation, theengineers work closely with the technicians in or-der to give support to some deviations which canhappen, so that the installation does not stop. Inthe meanwhile, the engineering work also continuesin the office by producing and finalizing the certifi-cation documents which prove that all the require-ments are complied and also will allow the aircraftto perform a flight test to be conducted some ex-perimental tests.

To solve deviations there are three solutions. Per-forming a drawing revision, prepare a deviationsheet or perform a concession are the three ac-tual solutions. The first is normally used when thedrawing will be reused more times. The second isthe fastest to perform, however, the process eachtime that the respective drawing is used must berepeated. The last happens when there is a devia-tion but this is not reported to engineering duringthe installation process and a document is preparedby the technicians department to be approved afterby engineering before the aircraft leaves the com-pany. In order to reduce the number of deviationsis provide a flexible design, where the drawing isdone in a way which can absorb the deviations ofthe aircraft.

In terms of certification, one option to show com-pliance with a specific requirement is with experi-mental results. This method is used by Jet Aviationto show compliance with the vibration and buffetingrequirement for the ka-band system where a flighttest is performed. In this test are installed sensorsto measure the acceleration in all the directions inthe frames where is installed the new system. Forthe A319 aircraft, object of this work , the test wasalso performed and the results show that no par-ticular frequency is subject to peakes of vibrationenergy indicating resonance or flutter for the rangeof frequencies of interest, below the 100Hz. Thus,the installation complies with the vibration and buf-feting requirements for certification.

The last stage, before apply for the STC is to pre-pare and conclude all the required documents whichwill prove that all the certification requirements arebeing fulfilled. Therefore, the required documentsare presented in the following list.

• CAF - Classification Assessment & ApplicationForm

• MDL - Master Data List

• CCS - Certification Compliance Sheet

• DLM - Drawing List Mechanical

• DLE - Drawing List Electrical

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• EIL - Electrical Item List

• ENO - Engineering Order

• WBS - Weight and Balance Statement

• SSR - Structural Substantiation Report

• DTA - Damage Tolerance Analysis

• ANA - Analysis Report

• DPS - subcontractor Document Process Slip

• FTP/R - Flight Test Plan / Report

• ICA - Instructions for Continued Airworthiness

A description of all these documents is given thethe complete document.

7. ConclusionsSeveral achievements have been fulfilled in thiswork, starting with the demonstration of the ben-efits for the industry by using a standard as theARINC 791 for the ka-band antenna, which is thebase of the design developed by Jet Aviation.

Secondly, the illustration of all the steps to certifya major modification as the installation of the ka-band system was accomplished, by going throughthe selection of the antenna location, the certifica-tion requirements, the mechanical design develop-ment, the installation and respective deviations, theflight test and the documentation required for theSupplemental Type Certificate (STC) application.

Thirdly, the major achievement in this work wasthe results obtained from the static and fatigueanalysis for the seven hypotheses in the parametricoptimization. The results proved that all designsare strong enough to withstand the selected criti-cal load case, rapid decompression of the radome.All the structural components presented a safetymargin higher than zero. In terms of fatigue, thealterations introduced in the fittings, doubler andintercostals led to a reduction on the fatigue life forall the hypothesis, although without any negativeimpact on the Airbus threshold. As final result,hypothesis number six - design with modified dou-blers and intercostals - was considered the optimaldue to the high percentage of weight reduction andlow reduction on the fatigue life number of cycles.To sum up, this initial parametric optimization wasuseful to find an optimal design which can be an op-tion in the future. However, further analysis mustbe conducted to verify the structural integrity forother load cases and to determinate the thresholdand intervals of inspection in order to be certify adesign.

Lastly, the important role played by the experi-mental flight test in the certification of the initialdesign of Jet Aviation was shown. The obtained

results during the flight test covering the vibrationand buffeting requirement proved that any peaks ofvibration energy indicating resonance or flutter oc-curs within the required range of frequencies. Withthese results plus the numerical ones was achieveda certified solution, where an EASA Supplemen-tal Type Certificate was issued with the number10071445, named as ”KA Band Satcom System”applicable for the Airbus A319 and A320 series air-crafts on the rear fuselage.

As a future work, there are two identified op-portunities:

The first one is to continue the scope of thepresent work, where the static analyses for the re-maining loads cases could be performed in orderto prove that all the seven hypotheses are strongenough under all the load cases. In addiction, thedamage tolerance analysis could also be conductedin order to find the intervals of inspections requiredfor certification. With these two topics covered, theresults could show if all the hypothesis are certifi-able designs or not.

The second option is to take the initial design de-veloped at Jet Aviation and update it for the newstandard ARINC 792, where there were 7 fittingsinterfacing the antenna and the aircraft airframe,there would be only six fittings, the most backward,the number seven would be removed. An initialstudy could be performed to verify if the updateddesign is still feasible in terms of stress and certifi-cability. In case of a negative result, what changescould be performed in the design in order to makeit feasible.

AcknowledgementsI would like to start off by expressing my gratitudeto Professor Afzal Suleman. All the availability pro-vided during this stage to give the required guidancewas indispensable.

Secondly, I would like to thank Diogo Alvim, mysupervisor in Jet Aviation, for all the support pro-vided in this dissertation during my traineeship. Itwas immeasurable all the guidance and availabilityshared with me.

Thirdly, I would like to express my gratitude toJet Aviation for receiving me in this journey andproviding all the resources required to perform thisdissertation. Furthermore, thank you for all thetraining, knowledge shared with me and the confi-dence shown in my daily work in the beginning ofmy professional career.

Finally, I would like to thank my friends and fam-ily for always push me forward in my life.

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[2] Certification specifications and acceptablemeans of compliance for large aeroplanes cs-25.Technical report, European Aviation SafetyAgency, November 2018. Amendment 22.

[3] S. A. A320 family ka band satcom sys-tem. Structural Substantiation Report JBSL-M-01326SSR-01, Jet Aviation AG, July 2019.

[4] F. A. Administration. Metallic materials prop-erties development and standardization, July2016. MMPDS-11.

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