Project for an Aircraft Change

Goncalo Ribeiro Batalhagoncalo.batalha@tecnico.ulisboa.pt

Instituto Superior Tecnico, Lisboa, Portugal

December 2016

Abstract

This dissertation sought to give answer to an existing need in an helicopter operator company (HTAHelicopters) by developing and design of an equipment with characteristics to be installed in a helicopter.In addition to the project, was prepared all the necessary technical documentation to the project besubmitted for EASA approval which is the civil aviation authority in Europe.

Based on the airworthiness requirements imposed by EASA and the airworthiness limitations im-posed by the aircraft manufacturer, it was created a geometric model for the equipment and weredeveloped attachment systems. Using the finite element method, these were subjected to static andfatigue analysis in order to define the final measures and validate the model developed in accordancewith the factor of safety defined for the project.

Finally, it was studied the influence that the equipment will have on the helicopter in terms ofaerodynamics, and weight and balance.

The conclusions were that the equipment which was developed meets what was intended by thecompany and moreover, since it was created all the necessary documentation and have been met all therequirements and limitations imposed by EASA and by the aircraft manufacturer, the project is readyto be certified.Keywords: Helicopter, Platform, EASA, FEM, Aerodynamic Load, Center of Gravity

1. Introduction

HTA Helicopters (HTA) is an helicopter operatorcompany that performs different types of aerialwork including installation of high voltage powerlines. However, the work inherent to high volt-age power lines is not limited only to his instal-lation. These also include the maintenance/repairand other types of work, such as mounting sign-posting balls on the cables. Also in these worksthe helicopter is an very useful working tool whichcan be used and it becomes very important since itallows a very fast and efficient implementation. Sothese are the jobs that HTA want be able to do too.

Normally, to be able to perform this type of workwith helicopter is need to install equipment in it,which allows the transportation of an operator out-side the helicopter. Thus, not only the operator canget better access to the power lines, as well as thepilot can, with the helicopter make a safer approachto them since they do not need to be so dangerouslyclose of the power lines.

The use of helicopters to perform this type ofwork is not new and therefore there is already somecertified equipment that can serve this purpose.However, the HTA conducted market research andfound that, for the aircraft in which is intended to

install the equipment (Airbus Helicopters AS 355F1), there is no equipment for this purpose. There-fore, the company decided that the solution wouldbe the creation of the desired equipment project,which is the purpose of this document.

2. Background2.1. Aviation Legislation

In the aeronautic industry any changes/ modifica-tions that wish to make on an aircraft, requirescertification by EASA (European Aviation SafetyAgency). A change in an aircraft is called modi-fication and may have different classifications, de-pending on its features. When a new aircraft modelis approved by EASA, it is issued a “Type Cer-tificate” (TC), which is the document that legallyguarantees its airworthiness. The same is requestedwhen it comes to the certification of a change, how-ever in this case is issued and then approved an “Supplemental Type Certificate ’(STC), which is asimilar document, although in this manner to bealways associated with a TC.

Thus, in the case of this project a modification,not only have to be met all the requirements ofEASA in order to be issued the STC, but also haveto be subject to the limitations and aircraft char-acteristics where it will be applied. Once all these

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requirements / limitations depend on many factorsand issues that arise in parallel with the develop-ment of the project, also the analysis of the docu-ments containing such information is made through-out the project. During the development of theproject have to be analyzed and considered theEASA and aircraft documents simultaneously.

2.2. Yield and FatigueMetallic materials have a characteristic behavior de-fined by an elastic range in which, when requestedby a particular load, the deformation occurs onlyduring the application period eventually, after un-loading, return to the initial geometry. When thestresses exceed a limit (yield stress), the materialstarts a plastic behavior and in that part of the de-formation no longer returns to the initial state.

Typically, one of the design criteria used for siz-ing a component is that the yield strength of thematerial cannot be exceeded, although sometimesuse a similar approach but for the tension rupture.In this dissertation, will be considered the design tothe yield strength, with an appropriate safety fac-tor, using the calculation of the von Mises stress tothe equivalent stress (equation 1) [3].

Seq

n≥ 1√

2[(σx − σy)2 + (σy − σz)2 + (σz − σx)2+

+6(τ2xy + τ2yz + τ2zx)]1/2 (1)

In the above equation, n is the factor of safetyand Seq is the equivalent stress. The terms σx, σyand σz are the normal stresses and τxy, τyz and τzxare the shear stresses. In addition, it is assumedthat the stress state is in equilibrium and thereforeτxy = τyx, tauyz = τzy and tauzx = τxz.

Most mechanical engineering structures are notsubject to the application of constant efforts i.e.typically the loads are dynamic and associated withthese is fatigue failure. Fatigue is a phenomenon de-fined as weakening of the metal or due to repeateddynamic loads and is considered the major cause offailure of mechanical structures.

The most common way to analyze a mechanicalcomponent fatigue is through the S-N curves. Thesecurves are obtained cyclically carrying a sample of agiven material and are counted the number of cyclesuntil failure occurs. However, these curves are de-termined for tension average zero, which althoughit is actually a shaft in rotation to pure bending, forexample, may not occur in a generic structure forthe need to find a criterion that allows convertingthe alternating tension and average tension in onealternating tension. To do that there are severalcriteria that can be used as the Soderberg Good-man’s Modified or Gerber [3]. Since the criterion

of Modified Goodman, being more conservative, isused quite often, this criterion was also chosen to beapplied in this project. This criterion is translatedby the equation refeq:goodman [3], where σa is thealternated stress, σm is the average stress, Se is thestress fatigue limit and Sut is the tensile strength.

σaSe

+σmSut

=1

n(2)

2.3. Numerical AnalysisIn this project we used the SolidWorks software notonly as a CAD tool, but also as a tool to performsimulations through numerical analysis. To performthe numerical analysis, Solidworks refers to the Fi-nite Element Method (FEM) which is the standardmethod of engineering design analysis. The FEMis the division of the CAD model into smaller partsand simpler forms, called elements, which thus be-comes a complex problem into several simpler prob-lems are solved simultaneously. The set of all theseelements is a finite element mesh.

When one finite element mesh is generated, Solid-works creates 3D tetrahedral solid elements, trian-gular shell elements 2D and 1D beam elements. Thesolid elements are more suitable for thicker parts,the shell elements for the modeling of thin workpieces (metal sheets, tubes, etc.) and the beamelements modeling structural components (beams,trusses, etc.). In the particular case of this project,as can be noticed later in this dissertation, the typesof most suitable elements to make the numericalanalysis are solid and shell.

SolidWorks has the option to select between gen-erate a mesh draft quality (composed by lineartetrahedral solid elements or linear triangular shellelements) and a high-quality mesh (composed byparabolic tetrahedral solid elements or parabolictriangular shell elements). Because it has morenodes and consequently more degrees of freedom,the quality mesh requires more computational re-sources. Although it offers better results and be-cause of that, it was used this type of mesh.

In numerical analysis, a critical aspect to consideris the analysis of the numerical error. Since in mostengineering problems, there is no analytical solu-tion that describes the behavior of the componentsaccording to the set of differential equations gov-erning their behavior, it is necessary to use anothermethod.

One of the most usual methods, it is an iterativeprocess in which multiple meshes are tested startingfrom coarser mesh for finer meshes, and the varia-tions are calculated from the results of one iterationto the next. In this way, we will be carrying outsuccessive mesh refinements until obtained a con-vergence of results in order to establish an accuratelevel of acceptable results.

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3. Methods3.1. Certification Specifications

The first step which was assumed to beginning thedevelopment of this project was an previously anal-ysis to the document “ EASA Certification Specifi-cations for Small Rotorcraft - Amendment 3” (CS-27) [2] and this refers to the FAA (Federal AviationAdministration) document “FAA Advisory Circu-lar 27-1B - Change 2 ” (AC 27-1B) [1]. The firstone (CS-27) is the document that contains the air-worthiness requirements that a helicopter with aweight below 3175 kg must meet in order to becertified. The second one (AC 27-1B) is a docu-ment that complements the first one and containsadditional and additional specific information foreach one of the requirements. These documentswere carefully targeted and intensive analysis, re-view and fulfillment to ensure that in spite of beingto develop a project of a modification to the TCof this aircraft, this still ensures unequivocally thecontinuity of initial airworthiness which has allowedits approval and issuance of the TC by EASA.

3.2. Design and Material selection

The helicopter Airbus Helicopters AS 355 F1 isavailable in the passenger compartment floor at-tachment points 11 comprise rings (figura 1), andis defined that the same would be used to securethe equipment. In figure 1, the lines of five andfour attachment points are identified as front andrear rings, respectively, and will be from here thesedesignations used.

Figure 1: Helicopter tiedown points

The type of equipment developed was a platformin “ L ” shape in which one of the parts of it, isfitted on the helicopter cabin floor, in the area ofthe rear seats, and is fixed to the three existentcentral rings on cabin floor near to the rear bulk-head. Two other fixed points in the front existentcentral rings on cabin floor near to the pilot andco-pilot seats. Consequently, another part of the“L” is located outside the passenger compartmentand extended out to a front region, very close tothe right side of the helicopter and in line of pilotview. In addition, it was determined that the plat-form is manufactured in tubular steel, and the topwas covered with an expanded metal sheet, weldedto the tubular frame. Also welded to the frame wasa seat for the operator.

Operation of this equipment determine that thetwo doors on the right side of the helicopter, pi-lot door and rear door such the two rear seats andrespective belts on the right side of the helicoptermust be removed.

Despite all the rings have the same dimensions,the front rings are mounted on the cabin floor in adifferent manner then the rear rings, and it was nec-essary to develop two different fastening systems.The main objectives were to create functional fas-tening systems but also in addition to make surethat is easy to assemble and all the parts are consti-tuted by as many as possible standard components.

3.2.1 Tubular StructureIt was determined that the tube to be used for theframe plate would be of circular cross section andthe material would be the alloy steel AISI 4130 Nor-malized. Since this steel alloy is a material thatcomplies with the standard AISI (American Ironand Steel Institute), can guarantee the compliancewith the airworthiness requirement CS 27.603 [2][1]. Furthermore, it is an alloy with low carbon con-tent and became therefore relatively easy to weld.

The design of this project was based on an CADmodel of the helicopter, at real world scale, andmeasurements made on the aircraft itself. In figure2 are presented views of helicopter CAD model, oneof which is an overall view of the model (with thedoors and rear seats on the right side removed), andthe other is an overall view of the cabin whereina large part of the equipment has been removedto facilitate the viewing zone which is based theplatform.

(a) Perspective view

(b) Cabin top view

Figure 2: Helicopter CAD model

Once set, the type of section and material of thetube, the next step was the preparation of the de-sign for the tubular structure. This would consist ofouter tubes, which define the contour of the plat-

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form, and smaller inner tubes (and hence lighter)that reinforce the entire structure. The outer andinner tubes are highlighted in blue in figures 3 (a)and 3 (b), respectively.

(a) Outer tubes (b) Inner tubes

Figure 3: Types of tubes identification

It was defined initially that the outer tubes havea diameter of 2” (50,8 mm) and a thickness of0,095” (2,413 mm) and that the inner tubes havea diameter of 1,125 ” (28,575 mm) and a thickness0,058”(1,4732 mm). These values were not definedbased on any specific criteria, and were used onlyas a starting point for the project. The final di-mensions were defined only after conducting sev-eral simulations, which some results are discussedin section 4.

3.2.2 Rear Attachment SystemIn figure 4 (a) a general view is shown of the fasten-ing system developed and in figure 4 (b) the sameis represented, but in an exploded view with thevarious numbered components. In figure 4 (b), itcan be seen that the components A and B are theexisting ring and the part to which this is attached,which are available on the helicopter, only are rep-resented in the figure for better understanding ofthe system.

Part 1 is an hexagonal socked head screw ISO7380 - M12×16 - 10.9 [10] AISI 1045 alloy steel, andthis material was also chosen for the components 2,3 and 5.

Part 2 is a spring washer - M12 [10]. The use ofthis element aims to comply with the airworthinessrequirement CS 27.607 [2] [1], since this type washerensures that the screw has loosened.

The part 3 is a rectangular washer, which hasbeen developed with the aim of increasing the areaover which the pressure spring washer rests.

The part 4 is the part that is designed to makethe connection between the clamping system andthe tubular structure. This piece is welded to thetubular structure is alloy steel AISI 4130 normalizedand has a uniform thickness of 0,125” (3,175 mm).

The part 5 is also a steel alloy AISI 1045 washer,but with some particularities. The purpose of thiswasher is to compensate the offset in the ring height(part B) to the component 4 can stand on a flatsurface. Thus, the thickness of the washer corre-sponds to the difference of heights and moreover

(a) Perspective view

(b) Exploded view

Figure 4: Rear Attachment System

has a chamfer which is what allows it to be settledin the ring.

Finally, the component 6 is the part designed towork inside the ring and in the holes of the compo-nents 4 and 5. The base of the part, which also hasa chamfer, rests on component A. Thus, the ring isseated on the base part in horizontal position. Fur-thermore, this piece also has a threaded hole wherethe screw is tightened to fix all components.

3.2.3 Front Attachment System

In figure 5 (a) is shown in a general view the fasten-ing system developed for the front rings and in fig-ure 5 (b) is represented the same, but in an explodedview with the various numbered components. Giventhe figure 5 (b) component A corresponds to thering that is available in the helicopter cabin floor,and again is only represented in the figure for easierunderstanding of the system.

Part 1 is the connecting part between the pipestructure and the front attachment system. Thispiece is a rectangular section tube made of AISI4130 Normalized which side has 0,75” (19,05 mm)thickness is 0,049”(1,2446 mm) and one end thereofis welded to the tubular structure .

Part 2 is a plug also made in Normalized alloysteel AISI 4130 which is welded to the component1. The part has a threaded hole for the screw canbe tightened.

Part 3 is a washer in which (similar to the rear

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(a) Perspective view

(b) Exploded view

Figure 5: Front Attachment System

part 5 of the system) the thickness compensatesfor the difference in the ring height, and also hasa chamfer which allows it be pressed against thering.

Part 4 is a spring washer - M12 same as that usedin the rear attachment system.

Finally, part 5 is an hexagonal socket head screwISO 7380 Interior - M12×40 - 10.9 [10]. Also this isidentical to that is used on the rear system, exceptfor having a bigger length.

It should be remarkable that, as for the tubu-lar structure, also for fixing systems, all dimensionsof the components set forth in this phase of theproject are subject to be changed according to theresults that were obtained in simulations that willbe studied in section 4. In figure 6 is representedthe tubular struture with the attachment systemsinstalled on the helicopter.

Figure 6: Tubular structure and attachment sys-tems

3.3. Expanded Sheet Metal and Seat

As mentioned in section 3.2, one of the elementswhich belongs to the platform, is an expanded sheet

metal. This expanded metal sheet is welded to thetubular structure along its contour, since they arequite similar. The shape are quite different only inthe area of the left rear seats, where the expandedsheet metal is trimmed to obtain enough space toallow passage of the bottom part of the rear seats toits usual attachment points despite of the structureinstallation. It was also developed for the platforma seat for the operator made in 2 mm soft steelmetal sheet which is welded to the expanded sheetmetal as shown in figure 7.

(a) Perspective view

(b) Seat location top view

Figure 7: Full equipment installed in the helicopter

The mesh used was a diamond mesh L62 T3060mild steel [11], and from the data provided by thesupplier a maximum load capacity has been set forthe area occupied by the seat 128 kg.

The airworthiness requirement CS 27.609 [2] [1]states that the structural elements of the aircraftmust be protected against corrosion. In order tomeet this requirement all the ferrous metal compo-nents constituting this platform and sub assemblyparts are immersed in an anti-corrosion bath.

3.4. Factor of Safety

The airworthiness requirement CS 27.303 [2] [1] re-quires to be used a minimum factor of safety of1,5 for the project. However, it is admitted fromanother value if that is more conservative. In or-der to determine the most appropriate safety fac-tor for the project, we used the method of Pugsley[9]. Through this method we obtained a minimumsafety factor for the design ns = 2, 325. Since thisvalue is more conservative than the required valueby the airworthiness requirements, this was the se-curity coefficient adopted for the development ofthis project.

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4. Results and Discussion

In this section is indicated and discusses the resultsobtained from the static and dynamic simulationsin Solidworks. Based on the results obtained, itwas settled the final dimensions of the equipmentcomponents.

4.1. Finite Element Mesh

In section 2.3, the importance of numerical erroranalysis associated with the results obtained fromthe FEM was explained. Therefore, before proceed-ing to the static and fatigue analysis were performedfinite element mesh studies in order the error to beminimized.

In order to simplify the analysis and the meshstudies reduce the time calculation were analyzedseparately the tubular structure and fastening sys-tems. In each case, and in accordance with theexplained in the section ref section: analisenumer-ica, we used the type of finite element more suitablefor the different types of the parts. Thus, for thetubular structure were used shell elements and solidelements for fastening systems.

For the tubular structure, during the static simu-lations was admitted that the operator has a max-imum weight of 115 kg. The operator seat has anapproximately weight of 5 kg. The assembly opera-tor plus seat sums are a total of 120 kg; this is whythis value was choose and used for the load appliedto the structure. In fastening systems simulations,forces that were applied were those resulting fromsimulations of the tubular structure.

The refinement study for the FEM resulted that,for the attachment systems, the finite elements havea size of 0,75 mm. In the case of the tubular struc-ture it would have 5 mm, except for those that wereconsidered critical zones where the size of the ele-ments is 1,25 mm. The critical areas are highlightedin blue in figure 8.

Figure 8: Critical areas

In the case of the expanded sheet metal, asthe weight of the assembly “seat + operator” issmaller than the maximum load which was obtainedthrough the supplier data for the mesh (128 kg),there was no need to carry out an analysis.

4.2. Static and Fatigue AnalysisOnce the finite element meshes to be used in thesimulations are definite, the conditions to performthe static analysis and fatigue are achieved. Thuswill be fulfilled the airworthiness requirements CS27.305, CS 27.307 and CS 27.613 [2] [1] which re-quires to be performed testing or static simulations,to demonstrate that the structure does not sufferpermanent deformation when subjected to the max-imum load, and dynamic simulations to prove theirresistance to fatigue.

4.2.1 Tubular StructureThe results obtained from the static simulation con-figuration of the tubular structure that has beeninitially set in the section 3.2.1 has concluded theneed to increase the dimensions of the tubes (di-ameter and/or thickness) in order to obtain accept-able stress values and factors of safety consideringthe minimum obtained safety coefficient (0,812) wasmuch lower than the minimum required (2,325). Sothey were carried out various combinations of dif-ferent tubes simulations in order to find the mostappropriate setting. It turned to a final configura-tion in which the inner tubes have a diameter of1,25” (31,750 mm) and a thickness of 0,095” (2,413mm), and outer tubes have a diameter of 2,5” (63,5mm) and a thickness of the inner tubes (except thesegment highlighted in blue) in the figure 9. Oncethe most critical points were found in this area,this segment was admitted to a thickness of 0,250”(6,350 mm). With this configuration, the value ob-tained for a minimum safety factor is 2,638 and thisway being above the minimum required value anda very much lower weight for a circumstance of aconfiguration in which all the outer tubes have thisthickness.

Figure 9: Tube segment with greater thickness

However, as mentioned earlier in section 2.2, theleading cause of failure of the mechanical structuresis fatigue, so it is necessary to analyze the behaviorof this setting when exposed to a simulation withdynamic forces. For this, were introduced in Solid-works data of the S-N curve [3] for the Normalizedsteel AISI 4130 and performed a dynamic simula-tion was to 106 cycles (infinite life). The minimumsafety factor obtained for this simulation (2,345)was held in reserve above the minimum required,so this configuration is fully validated.

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4.2.2 Rear Attachment SystemWith the configuration that was initially set for therear attachment system, in section 3.2.2, during thestatic simulation has showed a critical area corre-sponding to the area of contact between the partthat is welded to the tubular structure and the ringwho shows a minimum safety factor too low. In fact,by during this simulation it was noticeable that thecontact between these two parts was not requiredand therefore, a modification was made in the com-ponent that is welded to the tubular structure. Theadjustment passed by the creation of a chamfer onthe edge located in this area and thus there is nolonger contact between the two parts (Figure 10).

(a) Original system (b) Modified system

Figure 10: Modification applied to the rear attach-ment system

So, to validate this new configuration of the rearattachment system, it was made a new static sim-ulation whose result for the minimum safety factorwas 4,265. Since this new value is over 2,325, notonly the system is validated for the static load aswell there is still some range for the dynamic simu-lation.

The parameters established for the dynamic sim-ulation of the rear fastening system were the identi-cal as the dynamic simulations of the tubular struc-ture, however in this case, as there are differentmaterial components, in addition to the S-N curvealready used previously, it was necessary to also in-troduce in the SolidWorks the curve data on S-N [5]for steel AISI 1045 dynamic simulation resulted ina minimum safety factor to fatigue of 2,557 which ishigher than the minimum ratio set for the projectand indicates that the rear attachment system isvery well dimensioned.

4.2.3 Front Attachment SystemThe performed Static simulation in the front fasten-ing system has resulted in a fairly high minimumsafety coefficient (2,841) and because of that it waspossible to move to the dynamic simulation.

As for the dynamic simulation of the rear attach-ment system, since the forward system is also com-posed of parts in Normalized AISI 4130 and AISI1045, was necessary to use the S-N curves for eachmaterial. In the results obtained in the dynamicsimulation, it can be appreciated that the front at-tachment system has a resistance permissible fa-tigue, once the minimum safety coefficient obtained

at the end of 106 cycles (2,420) is greater than 2,325.

5. Aerodynamic Effects, Weight and Balance5.1. Aerodynamic Effects

When the equipment is installed on the helicopterthere is a part of it that is in an external area.Therefore, this part is subjected to aerodynamicloads created by the flow generated by the mainrotor rotation and by the horizontal forward move-ment of the helicopter. Thus, it is important toknow these loads and what is the influence thoseloads may have on the helicopter’s behavior.

5.1.1 Main Rotor Flow

In order to simplify the analysis, it was consideredthat the part of the platform that is beyond the in-terior of the cabin has the behavior of a flat plate ofnegligible thickness corresponding to the blue areathat is shown in figure 11.

Figure 11: Flat plate

Once it was admitted that the flow caused bythe rotor is perpendicular to the plate, this is anextreme case of a no fuselage body [4] that createsa great drag force. The calculation of this force,was made from the equation 3 [8] where CD is theplate drag coefficient (CD = 1,28 [8]), D is the dragforce in N ρair is the specific mass of air in kg/m3

(ρair = 1,225 kg/m3), A is the flat plate area in m2

(A = 1, 154 m2) and V is the flow velocity in m/s.

CD =D

12ρairAV

2(3)

Thus, the only thing remaining to be able to cal-culate the platform drag force (D) is the velocity ofthe flow created by the main rotor. For this, wasadmitted maximum power situation under “FlightManual AS 355 F1” [6], what means Pmax = 328kW for a rotor speed of Ω = 394 rpm. Thus, onecan calculate the rotor induced velocity (vi) fromthe equation 4 [7], where P is the power in W andT is the propulsion generated by the rotor (assumedto be equal to the maximum take-off weight of the

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helicopter corresponding to a mass of 2400 kg [6])i.e. T = 2400× 9, 8 = 23520 N.

vi =P

T(4)

From this equation it follows that the rotor in-duced velocity is vi = 13, 9 m/s. However, the ob-jective is to know the air speed of the flow generatedby the rotor (w) which is equal to twice the rotorinduced velocity (equation 5 [7]).

w = 2vi (5)

Therefore, the speed of air is w = 27,8 m/s thatis equivalent to the flow velocity (V ) from equation3. Thus, replacing the values in this equation ispossible to determine that D = 699,2 N. Assumingthat force is being applied to the center of the plate,it can be noticed that as well as the weight of theequipment itself or the weight of the operator, thisforce creates a moment about the axis of symmetryof the helicopter. For a better understanding ofthe influence of the moment created by the dragforce on the aircraft will be the same by a simulatedmass body placed in the center of the flat plate.This body will have an equivalent weight to the dragforce and therefore its mass is m = 699, 2/9, 8 =71,35 kg. In this way, a moment that will contributeto the displacement of the center of gravity of theaircraft will be created. Therefore, the influence ofthis load will be analyzed in more detail in sectionweight and balance.

5.1.2 Horizontal Forward MovementAssuming the main rotor is parallel to the cabinfloor, when it is at horizontal forward motion, typ-ically the helicopter is not perfectly in the horizon-tal position. In fact the helicopter “nose” sinksslightly and form an angle α with the horizontalplane. Since the equipment is in a position parallelto the cabin floor and is fixed to it, the platformassume the same angle than the helicopter (Figure12).

Figure 12: Flat plate area equal to the platformprojected area

Assuming an angle α = 20o, it can be seen, infigure 12, that the projected area of the platformis again a flat plate perpendicular to the flow andtherefore can be used again the equation 3 to cal-culate the drag force. To do that it is necessary

determine the area of the plate (AP ) and the veloc-ity flow (V ). From figure 12 AP can be calculatedlike AP = 0, 727 × 1, 588 × sin 20o = 0,395 m2. Itwas admitted that the helicopter moves at maxi-mum speed permitted by the aircraft manufacturerto operate without two doors (70 knots [6]), and afront wind of 30 knots, therefore V = 70+30 = 100knots = 51, 4 m/s. Substituting these values intothe equation 3 and keeping the values previouslyused for the other variables, it follows that D = 818N.

Assuming that the drag force acts on the centerof the flat plate, since it is displaced relative to thehelicopter plane of symmetry, it generate a momentthat, as well as the torque that the main rotor ap-plies to the helicopter, tends to make the helicopterturn on itself. Therefore, it is important to calcu-late the propulsion needed to be generated by thetail rotor to compensate the moment created by thedrag force and realize if it is within the operatingrange of the tail rotor.

To determine the value of the moment createdby the drag force of the plate (MD) relative to themain rotor shaft, it is just multiply the force by thedistance from the plane of symmetry of the aircraftto the point where it is applied (center of the flatplate). It follows that MD = 818 × 1, 213 = 992Nm. Dividing this value by the distance betweenthe two rotors, follows that the propulsion requiredfor the tail rotor is TTR = 992/6, 120 = 162 N.

Once calculated the propulsion needed to offsetthe plate drag force, must now find a normal rangerotor operating and verify if the calculated propul-sion does not exceed it. To do that, were admit-ted situations of maximum continuous torque andmaximum power. From equation 6 [7], knowing thepower (P ) W and the rotor speed in rad/s (con-stant and equal to Ω = 394 rpm = 41, 3 rad/s) itis possible to obtain the main rotor torque (MR).Dividing this value by the distance between the tworotors, one it is obtained the corresponding tail ro-tor propulsion.

MR =P

Ω(6)

According to the “Flight Manual AS 355 F1”[6] at maximum continuous torque situation, it isnecessary 73% of the available power, ie, P =0, 73× 328 = 239, 44 kW = 239440 W. Substituingthis value in equation 6, follows that MR = 5798Nm. Dividing this value by the distance betweenthe two rotors, it is obtained the propulsion gen-erated by the tail rotor TTR = 5798/6, 120 = 947N. For maximum power (P = 328 kW) follows thatMRmax = 7942 Nm and TTR = 7942/6, 120 = 1298N.

Subtracting the values obtained for the tail rotor

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propulsion, it is possible to determine an operatingrange ∆TTR. So it follows that ∆TTR = 1298 −947 = 351 N. Since this value is greater than theprupulsion that the tail rotor needs to generate tooffset the momentum created by the drag force ofthe plate (162 N), it means that the helicopter isable to operate with equipment installed.

5.2. Weight and BalanceAirworthiness requirements CS 27.25 and CS 27.27[2] [1] indicate that the limits imposed by the air-craft manufacturer for weight and location of thecenter of gravity, can never be exceeded. There-fore, it is necessary to understand which are theselimits and verify in what conditions the aircraft canoperate, when the equipment is installed.

The location of the center of gravity of the air-craft is defined from perpendicular measurementsto three reference plans that are represented in fig-ure 13 [6]. In this figure it can be seen that the Zreference plane is parallel to the cabin floor and isat a distance of 2,60 m above it. The Y plane cor-responds to the plan of symmetry of the helicopterand perpendicular to the cabin floor. The X plan isperpendicular to the two mentioned above, and isat a distance of 3,40 m in front of the center of themain rotor.

Figure 13: Reference plans of the rotorcraft [6]

The limitations imposed on the location of thecenter of gravity are the distance to the X plan (lon-gitudinal center of gravity) and the distance to theY plan relative to the (lateral center of gravity).The limits for the location of the lateral center ofgravity are up to 0,09 m to the right and up to 0,16m to the left of the Y plan. The limits for the longi-tudinal center of gravity depend on the weight tooand are presented in figure 14.

To determine the location of the center of gravity,must be taken into account for each element that isthe configuration of the aircraft (including its emptyweight), the distances that each one is to the ref-erence plans, its mass and the moment created forto each plan (obtained multiplying the distance bythe mass). The location of longitudinal and lateral

Figure 14: Limits for the longitudinal center ofgravity location [6]

centers of gravity is obtained from division betweenthe moments in X and Y, respectively, and the totalmass.

In this case, for calculating the center of grav-ity, in addition to the empty weight of the aircraft,were admitted in the calculations a pilot, the fueltanks, the equipment, the aerodynamic load of themain rotor flow that was calculated in section 5.1(D = 699, 2 N which is equivalent to a mass ofm = 71, 35 kg), the operator, the left hand luggagecompartment and passengers in the rear left side.Some scenarios were created to calculate the centerof gravity and it was possible define some operatingconditions, such as:

Only with the equipment installed, without op-erator and a with a pilot of 80 kg, the helicoptercan operate with 100% of the autonomy;

If to the previous situation is added an operatorwith 80 kg, it will be necessary to place a loadof 92 kg in the luggage compartment and theautonomy is reduced to 88,9%;

with the same pilot and an operator with themaximum permitted weight (115 kg), it willrequire the carriage of a passenger with 80 kgon the left seat, a charge in the luggage of 90kg and a reduction of autonomy to 69,3%.

However, these are just some of the extreme sit-uations in which the equipment can be used, andthere are several possible configurations.

6. ConclusionsBased on the results obtained in this project, it maybe concluded that the equipment demonstrates ahigh resistance to static loads and ensures resistanceto fatigue equivalent an infinite life since it allowsa minimum of 106 loadings.

The performed analysis to the influences of theequipment on the aircraft aerodynamics, allowed toconclude that the helicopter is able to operate withthe equipment installed. From the analysis to theweight and center of gravity location, it was pos-sible to conclude that the operation can be done

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with several configurations with at least 69,3% ofthe autonomy.

AcknowledgementsThe author would like to thank professors FilipeCunha and Jose Pereira for all the help and support.Eng. Joao Ferreira must also be thanked for hishelp in understanding the aviation rules and for thegathered data.

References[1] F. A. Administration. Advisory Circular AC

27-1B, 2014. Change 6.

[2] E. A. S. Agency. Certification Specificationsfor Small Rotorcraft CS-27, 2012. Amendment3.

[3] R. G. Budynas and J. K. Nisbett. Shigley’sMechanical Engineering Design. McGraw-Hill,9th edition, 2011.

[4] V. de Brederode. Aerodinmica Incompressvel:Fundamentos. IST Press, 1st edition, 2014.

[5] eFunda. High-cycle fatigue. Consultado a02/08/2016.

[6] Eurocopter. Flight Manual AS 355 F1, 1983.Revision 7.

[7] J. G. Leishman. Principles of Helicopter Aero-dynamics. Cambridge University Press, 2ndedition, 2006.

[8] G. R. C. NASA. Shape effects on drag. Con-sultado a 25/08/2016.

[9] S. R. Schmid, B. J. Hamrock, and B. O. Ja-cobson. Fundamentals of Machine Elements.CRC Press, 3rd edition, 2013.

[10] A. Silva, C. T. Ribeiro, J. Dias, and L. Sousa.Desenho Tcnico Moderno. LIDEL-Edies Tcni-cas, Lda., 12th edition, 2004.

[11] L. SJMetal Distendido. Produtos metlicos.Consultado a 10/07/2016.

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