1. KENT STATE UNIVERSITY
AERN 45700/55700: AIRCRAFT DESIGN
INSTRUCTOR: D. BLAKE STRINGER, PH.D.
Spring 2015
The Flash
by
Kent Aerospace, Inc.
Kayla Grass
Matthew Gazella JohathanHerman
AlexanderFlock StevenJohns
ThomasSpisak Scott Konesky
ObedAsamoah FranklinCosta
Daniel Abbas NicholasBrown
Guojie Wang Di Xu
8. 1. The Flash
1.1. Description
The Flashis consideredtobe a new classof aircraft; a lightpersonal jet. The marketfor
thistype of product isexpandingandshouldyieldhighprofitsbeginninginthe thirdyearof
production. The Flashwill be marketedtosmall businesses,flight schoolsandthe government,
to name a few. Its DGEN 380 TurbofanenginesbyPrice Inductionmake thisaircraftunique in
the sense of savingthe consumerinfuel andmaintenance costsaswell asweight. The aircraft
was designedforthe purpose of Price Inductioncreatingamarketforthe sale of theirengines.
The nominal cruisingaltitude is18,000 feetPA and the aircraftis capable of carryingthree
passengersinadditiontothe pilot. Itsstate of the art avionicspackage will attractmany
customersandmake the pilot’sjobmucheasier.
1.2. Summary of Key Parameters
Wing Geometry Performance Parameters Basic Performance
Dimensions(L) 34' 5" Engine Type DGEN380 Max Airspeed 250 kcas
WingSpan 37.34 ft StaticThrust HP 580 Cruise Speed 230 kcas
WingChord 7.66 - 1.91 ft Thrust at 18,000 ft 340 Service Ceiling 25,000 ft PA
AspectRatio 7.8 SFC 0.26 Range 800 NM
WingSurface 178.76 ft² MGWTO 4897 lbf Endurance 3.16 hrs
WingLoading 25 lb/ft²
1.3. Configuration layout
9. 2. REQUIREMENTSANALYSIS
2.1. Requirements Summary
Baseduponcurrent socio-economicdrivers,the followingrequirementshave beendetermined:
- The designwill be 14 CFR Part 23 compliant.
- The designteamwill utilize Part21 Certificationprocedures.
- The aircraft will utilize afly-by-wire systemtoreduce weight.
- The DGEN 380 engine incorporatesaFADECsystemforreducedmaintenance
costs as well asan electricstarterforweightreduction.
- Multi-functional displays willbe usedinthe cockpitforexceptionalpilot
situational awareness.
- The aircraft will be capable of carrying3 passengersinadditiontothe single pilot.
- The overall designwillincorporate techniquestoenable stable handling.
- Aircraftskinmade of compositeswill furtherreduce weight.
- The aircraft will have arange of 800 nautical miles.
- The aircraft will be capable of shorttake-offsandlandings.
10. 2.2. Mission Profile
Above isthe missionprofile expectedof the Flashwiththe associatedfuel burns expected
for eachleg,or missionsegment. Leg1. will includeengine start,possible missionequipment
checksand take-off. Leg2. includesthe aircraftclimbingtoa cruisingaltitude of 18,000 feetPA,
however,itiscapable of reaching25,000 feetPA. Leg 3. is the cruise portionwhere withthe
climbanddescentandloiterportionswill allow the aircrafttocoverup to 800 nautical miles.
Leg 4. is the descentandloiterportionwhere the expectedloitertime is20minutes. Finally,leg
5. is the landing,taxi andshutdownportion.
2.3. Reference Design Concepts (Baselines)
Eclipse 400
Dimensions _ Performance _ Powerplant PW610
length 29ft cruise speed 380mph max thrust 900lbs
wingspan 36ft range 1445mi bypass ratio 1.83
height 8ft 10in
service
ceiling 41000ft
empty
weight 2480lbs
gross weight 4480lbs
12. 3. TECHNICALDESIGN
3.1. Reference Design Concepts (Baselines)
3.2. Sizing Methodology
We came uponthe aircraftsizingforthe wingspan,length,andheightjustbylookingat
otheraircraft of a similarcategorythathave successfullyflownandlookingatwhattheir
respective dimensionsare.Forthe size of aircraft we are promotinginthisproject,a wingspan
from37-40 feetseemedtobe whatall the successfullyflownvery-lightpersonal jetshave as
theirwingspan.The lengthwe came towasdue to the inspirationsmentionedearilerwithan
average lengthof 35-40 ft beingthe mostprevalent.Alsothe lengthwasinfluencedbythe
placementof the enginesaswe decidedearlyonforthe twinDGEN enginestobe mountedto
the side of the rearward fuselage.The heightwasinfluencedbyotheraircraftof the same class
as before,withfurtherinfluence bythe seatingarrangement.We neededtodecide where the
passengerswouldsitandhowtall anaverage personsittinginthe type of seatwe wantedwould
equate to.The other sizingparameterssuchasweightandrange were calculatedbythe class
13. individuallyandthe chosennumberstakenfromthose thatwere deemedmore accurate than
the rest.
3.2.1. Design Space
Since the aircraftwas designedaroundthe engines,we knew fromthe beginningwhatour
altitudesof operationwouldbe. Price Inductionhadalreadydeterminedthe enginestobe
operationallysounduptoan altitude of 25,000 feetPA. Consideringthe powerthe DGEN 380
produces,a lighterjetwasthe onlyviable option.
3.3. Assumptions
Major assumptionsaffectingthe design:
3.3.1. Assumptions Used for Lift-to-Drag Ratio (𝑳 𝑫⁄ )
𝐿
𝐷 𝑚𝑎𝑥
estimationconstant: 𝐾𝐿𝐷 =15.5 for civil Jets
Wettedarearatio: 𝑆 𝑤𝑒𝑡 𝑆 𝑟𝑒𝑓⁄ = 4.1
AspectRatio:AR= 7.8 for General Aviation-twinengine
3.3.2. Assumptions Used for Initial Sizing
Range:R = 800 [𝑚𝑛𝑖]
LoiterTime-Endurance:E= 20 [𝑚𝑖𝑛]
Cruise SpeedatFL180: 𝑀𝑐𝑟𝑢𝑖 𝑠 𝑒 = 𝑉𝑐𝑟𝑢𝑖𝑠𝑒 = 0.35 Mach
Constantinemptyweightfractionequation:A=1.51 for General aviation-twin engine
Constantinemptyweightfractionequation:C= -0.10 forGeneral aviation-twinengine
Variable sweptconstant: 𝐾𝑉𝑆 =1.00 forfixedsweep
14. 3.3.3. Assumptions Used for Thrust-to-Weight Ratio (𝑻 𝑾⁄ )
Maximumspeed: 𝑀 𝑚𝑎𝑥 = 1.2 𝑀𝑐𝑟𝑢𝑖𝑠𝑒
ConstantinT/W statistical estimationequation:a=0.267 forJet Transport
Constantin T/W statistical estimationequation:C=0.363 forJet Transport
3.3.4. Assumptions Used for Wing Loading (𝑾 𝑺⁄ )
Take off distance: 𝑆 𝑡 𝑜⁄ = 2500 [𝑓𝑡]
Take off Parameter: TOP = 120
ApproachSpeed: 𝑉𝐴𝑃𝐻 = 120 [𝑓𝑡]
OswaldEfficiency:e = 0.8
Zero-Left-Dragcoefficient: 𝐶 𝐷0
= 0.015 for Jets
3.3.5. Assumptions Used for Wing, Tail, and Fuselage Geometry
Taper Ratioof wing: 𝜆 𝑤 = 0.25
ConstantinFuselage lengthequation:a= 0.67 for Jettransport
ConstantinFuselage lengthequation:c= 0.43 for Jettransport
Taper Ratioof tails: 𝜆ℎ= 𝜆 𝑣= 𝜆 𝑤 = 0.25
Aspectratioof horizontal tail: 𝐴𝑅ℎ =2 3⁄ 𝐴𝑅
Aspectratioof vertical tail: 𝐴𝑅 𝑣 = 1.5
Horizontal tail volume coefficient: 𝑐 𝐻𝑇 =0.90 fortwinturboprop
15. Vertical tail volume coefficient: 𝑐 𝑉𝑇 =0.08 for twinturboprop
3.4. Wing and Tail Geometry
Thissectiondiscussesthe airfoil selectionandparametersfor geometry sizing of wings,
tails and fuselage.
3.4.1. Airfoil Selection
The selection of airfoil is one of the most critical phases in the conceptual design. The
characteristicsof a specificairfoil will have asignificanteffectonthe performance of wings. The
ideal selectionisthe airfoil whichiscapable of producinghighliftandlow drag. Airfoil selection
largely depends on the general considerations of the following factors:
- Airfoil geometry,suchascamberand thickness;
- Aerodynamiccharacteristics,suchasliftanddrag characteristics;
- Stall characteristics;
- Otherconsiderations,suchasReynoldsnumber,structural layout,anddifferent
components(Raymer,2012).
A varietyof airfoilshave beendevelopedbydifferentinstitutions.Inthe selectionof this
design, the consideration will only depend on the airfoils developed by NACA. Four series of
airfoilsdeveloped by NACA are widely used in modern aircraft, the four-digit series, five-digit
series, the six-series airfoils, and seven-series airfoils. By comparing several airfoils from the
above factors,it isdesirable toselectthe airfoil commensurate tothe ideal one.However,there
are always some tradeoffs through the process of selecting.
3.4.2 Wing Geometry
Basedon the TOGW determinedatthe initial sizing,the coefficientliftof the ideal airfoil
during cruise is determined with the ideal coefficient lift (𝐶𝑙 𝑖𝑑𝑒𝑎𝑙
)to be 0.18. The design lift of
16. coefficientis1.11 whichisthe liftcoefficient( 𝐶𝑙 𝑐𝑟𝑢𝑖𝑠𝑒
) associated to the (𝐿 𝐷) 𝑚𝑎𝑥⁄ . In addition,
other considerations should be included in the selection of tip airfoil. The report Summary of
Airfoil Datapublishedbythe National AdvisoryCommittee forAeronautic (1945) states that it is
desirable for tip selection to have a high maximum lift coefficient ( 𝐶𝑙 𝑚𝑎𝑥
) and a large Critical
angle of attack (∝ 𝑠𝑡𝑎𝑙𝑙) in orderto increase the stall performance (NACA,1945).Asfor thickness,
the thicker the airfoil is, the more lift the airfoil will produce. Consequently, selecting the
thickest airfoil is advantageous.
Takingall above requirementsintoconsideration,the criteriain response to the pivotal
factors for airfoil selections are listed below:
1. Maximumliftcoefficient(𝐶𝑙 𝑚𝑎𝑥
) isthe highest.
2. Critical angle of attack (∝ 𝑠𝑡𝑎𝑙𝑙) isthe highest.
3. Coefficientof pitchingmoment(𝐶 𝑚) isclose to0
4. Maximumlift-to-dragratio(𝐶𝑙 𝐶 𝑑⁄ 𝑚𝑎𝑥) at cruise isclose to (𝐿 𝐷) 𝑚𝑎𝑥⁄
5. Liftcoefficient(𝐶𝑙) of maximumlift-to-dragratio(𝐶𝑙 𝐶 𝑑 𝑚𝑎𝑥⁄ ) atcruise isclose to 𝐶𝑙 𝑖𝑑𝑒𝑎𝑙
6. MinimumDrag coefficient(𝐶 𝑑 𝑚𝑖𝑛
) isthe lowest
7. Liftcoefficient(𝐶𝑙) of minimumdragcoefficient(𝐶 𝑑 𝑚𝑖𝑛
) atcruise isclose to 𝐶𝑙 𝑐𝑟𝑢𝑖𝑠𝑒
8. Thicknessratio(𝑡 𝑐⁄ ) ishighest
After comparing eleven airfoils listed in appendix 3-2, each airfoil is rated from the above
eight criteria. The airfoils with the highest rates are NACA 23012 and NACA 654-221. With
furtherconsiderationsonthe thicknessforrootandtip selections, the thickness of root section
ispreferable to be thick to provide space for fuel and equipment (Abbott, Doenhoff & Stivers,
1945). According to Dr. Sadraey (2012) in his book Aircraft Design: A System Engineering
Approach, “As a guidance; the typical values for the airfoil maximum thickness-to-chord ratio
(𝑡 𝑐⁄ ) of majorityof aircraftare about6% to 18%.” (Sadraey,2012). For differenttypesof aircraft
17. in regard to speed, the maximum 𝑡 𝑐⁄ is between 9% to 12% for a high subsonic passenger
aircraft, and 15% to 18% for a low speed, high lift requiring aircraft (Sadraey, 2012). Therefore,
to optimize the airfoil to promote the performance of the designed aircraft, the airfoil NACA
23012 isselectedforthe tipsof the wings with NACA 23015 for the roots. With the selection of
those two airfoil, the maximum coefficient of lift for wings ( 𝐶 𝐿 𝑚𝑎𝑥
) is determined to be 1.55.
In addition to airfoil selection, there are other key factors regarding the aircraft
wings. Wing location, wing area, wingspan, and sweep angle have major effects on overall
aircraft performance.
For thisaircraft,a lowwinghasbeenselected.Whilebothhigh wing and low wing have
benefits, low wing is usually preferred for training purposes, and is also commonly found on
mostjetaircraft. Lowwingofferseasieraccesstofueling. Low wing also allows easier access to
the enginesformaintenance purposes,andallowsthe student to easier be able to monitor the
engine duringflight.The easieraccessengines also help reduce maintenance costs in the long-
run, withshorterinspectiontimes.Low wingalsooffersbettervisibilityduringturningandother
aerial maneuvers.Stowing landinggearsispossible forbothhighwingandlow wingaircraft, but
is much easier in low wing aircraft, as the structure is much more available to the gear.
To determine the total wingarearequired,itisnecessarytouse the following equation
(E.q.3.5.2-1). The calculated weight used in the equation is 4,897 pounds. The wing loading
calculation used is 27.3938. The resulting wing planform area comes out to be 178.76 square
feet.
𝑆 = 𝑊 / (𝑊 𝑆)⁄ ------------------------------E.q.3.5.2-1
A total wingspan of 37.34 feet has been calculated. In order to calculate wingspan, the
aspectratio isassumedas7.8 for thiscalculation.The formula(E.q.3.5.2-2) explainedisshownas
follows.
18. 𝑏 = √ 𝐴 ∗ 𝑆 --------------------------------------3.5.2-1
Sweep angle is another important parameter regarding wing design. Changing the
sweepangle hasmanyeffects on performance, such as stability due to shifting the MAC of the
wing, or helping to avoid the onset of shock waves. From historical statistics (Raymer, 2012), a
sweep angle of approximately 2.0 degrees would be sufficient for the given aircraft.
3.4.3 Fuselage Geometry
The layoutof a fuselage isgenerallydependentonthe TOGW and the functionof the
aircraft.The primaryfunctionof the designedaircraftistocarry passengers.Giventhe number
of passengersandcrews,the lengthanddiameterof the fuselage will eventuallybe determined.
However,since the proposedaircraftisalsodesignedtoundertake someothertasksmore than
carryingpassengers,otherconsiderationsshouldalsobe takenintoaccount.Fromthe historical
statistics,the followingequation(Eq.3.5.3-1) will be usedtodetermine the lengthof the
fuselage:
𝐿 𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒 = 𝑎𝑊0
𝐶
---------------------------------------Eq.3.5.3-1
The TOGW has been determined. Based on the major assumptions made in section 3.4.5,
the length of the fuselage is calculated to be 25.87 [𝑓𝑡]. The maximum fuselage diameter is
determined by the ratio between fuselage length and maximum fuselage diameter, which is
referredasfinenessratio.Tominimize the drag produced by the fuselage, the fineness ratio is
around 3 (Raymer, 2012). As a result, the maximum diameter of the fuselage is set to be 8.62
[𝑓𝑡].
19. 3.4.4 Tail Geometry
The major functionof the horizontal tail istocreate a nose upmomentto counter
the nose-downmomentcreatedbythe wings.Whenthe elevatororrudderisnot deployed,the
tail isexpectedtoproduce zeroor little amountof liftormoment.Toachieve these two
purposes,symmetricairfoils are suitable selections. Toseveral general aviationaircraft,the
NACA 0012 andthe NACA0009 are appliedfortails. Additionally,outof the considerationfor
compressibilityeffect,the tails’thicknessshouldbe lessthanthe thicknessof the wings
(Sadraey,2012). Giventhe reasonsabove,the tail airfoil forthe new designischosentobe
NACA 0009.
The configurationof a tail isinfluencedbytrimming,stability,controllability,
operational requirements,airworthinessandsome otherlimits.Toproperlyapplythe
configurationof atail requiresprofessional analysisonthe above factors.MostGA aircraft and
airline aircraftuse conventional tail becauseitprovidessome benefitssuchaslightweight,
efficient,andperformsatregularflightconditions(Raymer,2012).With limitedbudgetsand
manufacturinglevel,the conventional tail will be employedinthe designedaircraft.
The geometryof a tail is determinedbyitsprimaryfunction.The tail geometryisdirectly
relatedtothe winggeometry. Besides,the tail size isalsorelatedtothe lengthof the fuselage
and the positionof the engines.The tail armis about50% to 55% of the fuselage lengthforan
aircraft withthe enginesonthe wings,about45% to50% foraft-mountedengines(Raymer,
2012). With respecttothe drawingof the new design,the enginesof proposedaircraftare
mountedonthe side of aft-fuselage.Therefore,the armlengthsof horizontal tail (𝐿 𝐻𝑇) and
vertical tail (𝐿 𝑉𝑇)are decidedtobe 50% of the fuselage length(𝐿 𝑓𝑢𝑠𝑒𝑙𝑎𝑔𝑒).
20. The methodof calculatingthe parametersof tailsispertainingtothe definitionof tail
volume coefficient.The followingtwoequations,Eq.3.5.4-1and Eq. 3.5.4-2 define the horizontal
tail volume coefficientandthe vertical tail volumecoefficientrespectively:
𝑐 𝐻𝑇 =
𝐿 𝐻𝑇 𝑆 𝐻𝑇
𝐶 𝑊 𝑆 𝑊
----------------------------------Eq.3.5.4-1
𝑐 𝑉𝑇 =
𝐿 𝑉𝑇 𝑆 𝑉𝑇
𝑏 𝑊 𝑆 𝑊
-----------------------------------Eq.3.5.4-2
Withthe resultsof winggeometrycalculationsandassumptionsmade insection
3.4.5, the area of the horizontal tail is59.55 [𝑓𝑡2] and the area of the vertical tail is41.30 [𝑓𝑡2].
The methodto calculate the othertail parameters,suchasroot chord, tipchord,span, lengthof
the MAC, and locationof the AC isthe same as the methodusedforwinggeometrycalculation.
Those parameterswill be listedundersection3.11.3.
3.5. Thrust-to-Weight Ratio
The wingswere primarilydesignedtosupportstabile handlingandlongendurance applications.
The thrust to weightratioiscalculatedtobe 25 lb/ft².
3.6. Introduction to Powerplant Data
Performance isone of the mostsoughtafterfactors whendevelopinganew aircraft.It does
not matterwhatkindof aircraft:helicopter,airplane,military,transport,orcargo.You will
alwaysrelyonthe performance of the aircraftto complete the taskat hand.The
mission/objective couldbe takingpassengersfromChicagotoNew York,or a Militaryjointstrike
fighterneedingtotake off froma carrier to dropa payloadovera conflictzone inanother
country.Each missionhasitsown setof establishedperformanceparametersthatthe aircraft
needstomeetinorderto successfullycompletethe objective.Wheninthe designphase,itis
21. necessarytolisteachmissionthe aircraftbeingdesignedneedstocomplete soyoucanstart to
analyze whatkindof performance will have tobe met.
3.6.1. Introduction to The Flash
Ourteam of engineersanddesignersatKentState hadto immediatelyaddressthe
performance factorsforour aircraft.Thisis because we hadto designthe entire plane arounda
DGEN 380 turbine engine.Thisturbofanengine hasalreadybeendesigned,developed,andhas
beguntestingtoconfirmitsairworthiness.Afterithasbeencertified,itisthenreadytomove
ontothe productionphase.We are workingwithPrice Inductiontohelpdesignanddevelopa
lightpersonal jetthatwill revolutionize the lightjetindustry.Price Inductionhasalsodeveloped
the SolutionsWesttCS/BV DGEN 380 engine simulator.Thispieceof technologyisatestengine
benchwhere the usercan recordand analyze differentparametersoccurringinsidethe engine
to gaina betterunderstandingof itspropulsionproperties.Thisalsogivesthe usersthe
possibilitytodesignanentire airplane aroundthistestenginebench.Thisgave ourteamat Kent
State an advantage because we were able tosimulate whatkindof performanceparametersthe
engineswillbe exposedtowhilethe aircraftiscompletingitsintendedmission.
3.6.2. Flash Performance
We are workingtogetherwithPrice Inductiononthisprojectof designinganew personal
lightjet,sowe alreadyknewwhatkindof engineswe wouldbe using.Theirengineershave
createda highlyefficientenginethathasa bypassratio of 7.6, and isverylightweightcomingin
at only175 pounds.Thisengine utilizesaFull AuthorityDigital EngineControl orFADECforits
powermanagement.AccordingtoPrice Induction,anall-electricconcepthasbeenvalidated,
such as an electricstarterandignitionsystem.Thisisverycritical because the onboard
generatoriscapable of producing6 kw of powerwhere 1.5 is neededforenginecomponents
22. and 4.5 can be usedfor variousairframe systemssuchasavionics,hydraulicsetc.Theyhave
createdthisengine torevolutionize the personal lightjetmarket.
The turbofanengine wasourbiggestsingle limitingfactorwhendesigningthe flash,aswe
had to come up withan aircraftdesignthatwouldperfectlyfittheseengine’sperformance and
characteristics.Whencreatingthe chartsand graphs of the performance we splititupintotwo
categorieseachwiththeirowndistinctparameterswe settocovera broadrange of scenarios
that the Flashwouldbe exposedtoina normal missionprofile.ForStaticperformance,we
simulatedthe enginestoexperience differentaltitudeswiththe correspondingstandard
temperatures,howeverwe didnotmeasure the performanceswiththe velocityof the airplane.
We calculatedthe staticperformance atboth100% throttle fora take-off condition,andalsoat
43% throttle forcruise conditions.Non-Staticperformances,againwe simulatedthe engines
performance atthe same altitudes,butincludedthe velocityof the aircraftinthe setof
parameterswe were able tochange on the testengine bench.Fornon-staticwe performedthe
take-off,cruise andalsomax speedconditions.
There were alsosome discrepanciesinourdata researchthat we found.One of the
challengeswe were facedwithwasdealingwithasimulatorthatonlyrecordedthe dataand
parametersusingthe metricsystem.Before we couldstartanalyzingthe datawe collected,we
had to convertthrust,fuel consumption,specificfuel consumptionandotherparametersthat
we collectedtothe Englishsystem.The biggesthurdle thatcame aboutwaswhenwe were
lookingatthe specificfuel consumption.AfterconvertingfromkgFuel/kgThrust/hrto
lbFuel/lbThrust/hr.We thenbegantolookata calculatedversionof specificfuelconsumption
usingthe equationFuel consumption/Total thrusttosee how thisdatacompared.I foundthese
numberstobe completelydifferent.We are still unsure of whetherthisisanissue withthe
simulatororwiththe data we collectedtoinputintothatequation.
23. Withinthe pastweek,ourteamwasable to complete afinished3Dmodel of the Flashand
was able toput it inKentStateswindtunnel.Fromthe datawe collectedfordragforces,we are
able to come up withan estimatedthrustrequiredcurve thatwill be requiredforouraircraftat
differentvelocitiesandaltitudes.Again,Iwouldlike tostressthatthisisjustan estimate
because ourmodel didnothave the smoothestsurface,whichwilladdtothe parasiticdrag of
the 3D printedaircraft.Alsobecause thisisascale model thathas a scale of 1:58 inchesitis very
hard to preciselycalculatehowthe actual airplane willperform. There are manydifferent
factors thatgo intoeach testfor scale modelsandmaynot be the same factorsor conditions
testingactual lightjetaircraft.Anotherexample of thiscouldbe thatour 3D printedmodel has
fullycoveredenginesthatdonot allow airto flow throughthem.Thiswill greatlyincrease the
drag of the Aircraft.
3.6.3. DGEN 380 Specifications and Performance
Condition Thrust SpecificFuel Consumption
Thrust AtTake Off power(SLS,Mach: 0) 570 lbf 0.44
Thrust at Max Continuous(FL100,Mach: 0.338) 240 lbf 0.78
Thrust at Max Continuous(FL180,Mach: 0.4) 185 lbf 0.80
-Table 3.1
*These are performancesforonlyone DGEN 380 TurbofanEngine.
Price Inductionhascome up with2 standard applicationsforthisenginelistedbelow.
StandardApplications: 2 Seats(Single Engine) 4+1 Seats(Multi Engine)
Max Take off Weight: 1,980 lb 3,640 lb
WingLoading: 25 lb/ft^2 25 lb/ft^2
Entire Surface area of A/c: 380 ft^2 700 ft^2
Max Cruise Airspeed: 247 mph 288 mph
24. Take Off Distance: 1,575 ft 1,900 ft
Fuel Onboard: 550 lb 1,050 lb
Range at Cruise (FL120) 615 Nm+ 45 minutes 600 Nm+ 45 min
Range at Cruise (FL220) 810 Nm+ 45 minutes 800 Nm+ 45 min
3.6.4. Static Takeoff Condition 0-15,000 Feet
29. Afteranalyzingthe dataforour cruise condition,we have come toa conclusiononwhythere is
a significantincrease inboththrustandfuel consumption. We believethatthisisbecause Price
Inductionhasdesignedthisaircrafttobe at optional performance ataround12,000-16,000 feet.
Thischaracteristicisalsoprevalentinsome of the otherconditionsthe enginewasexposedto.
31. 3.6.9. Wind Tunnel Data and L/D Curve
Creatinga 3D printedmodel of the aircraftwe designedgave usa muchbetterunderstanding
of howour aircraftwill actuallyperforminreal life conditions.Variousdatawascollectedin
preparationtocreate a Lift/Dragcurve more commonlyknownasthe thrust requiredcurve.
Furtheranalysiswasperformedtocalculate the coefficientof dragforthe 3D model.Thisisjust
an estimate,andmaynotbe quite ashighof a numberonthe real Flashafteritis certifiedand
produced.These calculationswere alsoperformedatSLSconditionswiththe airdensitybeing
0.00237 slugs/ft3
.The higherVelocitiescreatedmore accurate coefficientsof drag,sothe main
focuswill be onthose numbers.
32. Thisis our 3D printedmodel before itwassandedmade smooth. AdamZuckermanand
some membersof ouraircraft designteamspentcountlesshourstoperfectthe surface of our
model inorderto getit readyfor the windtunnel.Thiswasdone tolessenthe parasiticdrag
that will be producedfromroughsurfaces.Asyouwill see below intable 3.4,our parasiticdrag
was incrediblyhigh.Thisledtoaveryhighthrust-requiredneededtoovercome thisdrag.
34.
FD DragForce
AirDensity 0.00237(slugs/ ft3
)
V Velocity V(Fps)
A PlanformArea 0.0523ft2
CD
2 0.01102
0.00237 (26.72
) 0.0523
CD 0.249
CD
2 0.01543
0.00237 (34.82
) 0.0523
CD 0.205
CD
2 0.01764
0.00237 (402
) 0.0523
CD 0.1779
CD
2 0.01984
0.00237 (43.32
) 0.0523
CD 0.171
CD
2 0.02425
0.00237 (46.62
) 0.0523
CD 0.1797
CD
2 0.02866
0.00237 (502
) 0.0523
CD 0.185
The averagesof these drag coefficientsare whatwill be usedwhencreatingthe thrust-required
curve for the 3D printedmodel.Againitisimportanttonote thatthese characteristicswill vary
for the actual aircraft since some estimationwasinvolvedinthe process.
CD
2 FD
V 2
A
35.
CDA
0.249 0.205 0.1779 0.171 0.1797 0.185
6
CDA
0.1946
From thisaverage dragcoefficient,we are now able tocalculate the parasiticand
induceddragproducedbyour aircraft.This will thenbe usedtocalculate the thrustthat is
requiredtoovercome thisdraginsteadylevel flight.
Altitude
Density
(rho) S Weight
Oswald's
e
Aspect
Ratio K pi CD
MSL 0.00237 178.76
4897
lbf 0.8 7.8 0.05101108 3.1416 0.1946
From thiscalculateddata,we can now create a thrust-requiredcurve thatouraircraft will need
to meetforsteadylevel flight.
36. 3.7. Wing Loading Data
The wingloading, 𝑊 𝑆⁄ isthe ratioof weighttothe wingreference area.Certain
performancesof anaircraft,as stall speed,rate of climb,takeoff andlandingdistance,lift
producedbywings,etc.are affectedbywingloading.Todetermine the wingloadingfor
designedaircraft,the wingloadingmustbe comparedatsome commonconditions.The
followingsectionswillpresentthe discussiononthe calculationsof wingloadingatthree
differentconditions,stall,takeoff,andcruise.
3.7.1. Stall
The stall speedof an aircraftis directlydeterminedbythe wingloadingandmaximum
liftcoefficient.Stallspeedisone of the majorsafetyfactorsthat needtobe paidspecial
attentiontoinaviation.Several fatal accidentsoccurannuallydue tofailure tomaintainflying
speed.Todetermine the wingloadingrequiredtomeetacertainstall speed,liftmustequal
weight.Derivedfromthe liftequationatstall condition( E.q.3.8.1-1), the wingloading
requirementcanbe determined.
𝑊 = 𝐿 = 𝑞 𝑠𝑡𝑎𝑙𝑙 𝑆𝐶𝐿 𝑚𝑎𝑥
=
1
2
𝜌0 𝑉𝑠𝑡𝑎𝑙𝑙
2
𝐶 𝐿 𝑚𝑎𝑥
----------------E.q.3.8.1-1
37. The formulafor wingloadingrequirementforstall givesaresultof 44.79 [𝑙𝑏𝑓 𝑓𝑡2]⁄ . This
calculationisalsodone with 𝐶 𝐿 𝑚𝑎𝑥
of 1.55, a stall velocityof 155.83 fps,and air densityof
0.0024[𝑠𝑙𝑢𝑔 𝑓𝑡3]⁄ at sea level standard(SLS).
3.7.2. Takeoff
To determine the requiredwingloadingtomeetagiventakeoff distance requirement,
the followingexpression(E.q.3.8.2-1) isused.Inthiscalculation,the assumedtakeoff distance is
2,500 feet.The takeoff parameter(TOP) canbe foundfromfig5.4 inthe Raymertext,Aircraft
Design:A Conceptual Approach(Raymer,2012).
𝑊 𝑆⁄ = (𝑇𝑂𝑃)𝜎𝐶 𝐿 𝑇/𝑂
(𝑇 𝑊⁄ ) 𝑇/𝑂 --------------------E.q.3.8.2-1
The wingloadingrequirementfortakeoff comesouttobe 29.96[𝑙𝑏𝑓 𝑓𝑡2]⁄ . The calculated
𝐶 𝐿 𝑇/𝑂
is 1.281. Othervariablesusedinthe equationare the TOPwhichisassumedtobe 120,
densityratioof 1, and (𝑇 𝑊⁄ ) 𝑇/𝑂 of 0.1949.
3.7.3. Cruise
Determiningawingloadingforcruise is utmostimportant.The cruise conditionis
typicallythe mostdesignedaroundfactoronan aircraft. Choosingawingloadingfactorthat
directlysuitsthe cruise conditionforamaximumrange isproblematic.The wingloadingfactor
for a maximumrange ismuchhigherthan the wingloadingfactorrequiredforstall andother
characteristics.Itwouldbe unsafe toflywithsucha small wing,hence where understandingthe
importance of trade-offscomesintoplay.Tocalculate the wingloadingformaximumrange, the
followingequation(E.q.3.8.3-1) istobe used.
𝑊 𝑆⁄ = 𝑞 √ 𝜋𝐴𝑒𝐶 𝐷0
/3---------------------------E.q.3.8.3-1
The dynamicpressat the cruise conditionisdeterminebythe airdensityatcruise
altitude (FLl80) andthe cruise speed.Forjetaircraft, the Oswaldefficiency(e) andthe zero-lift
38. drag coefficient(𝐶 𝐷0
) are statisticallyassumedtobe 0.8 and 0.015 respectively.Aftertakingall
the variantsintothe above formula,the wingloadingatthe cruise conditioniscalculatedtobe
27.39[𝑙𝑏𝑓 𝑓𝑡2]⁄ .
3.7.4. Discussionof the Wing Loading
To determine the endwingloadingrequirement,all differentflightoperationsmustbe
considered,suchasstall,landing,takeoff,andcruise.Topickthe exactwingloadingthatwill be
usedforthe design process,the lowestcalculatedfromall of the flightconditionsistobe used.
Afterthe comparingthe resultsof the above calculation,the requiredwingloadingis
27.39[𝑙𝑏𝑓 𝑓𝑡2]⁄ . Selectingthe lowestwingloadingimpliesthatthe aircrafthas enoughliftbeing
producedbythe wing,for the givenweight.
3.8. Sizing Results and Design Selection
3.8.1. Sizing Variability and Optimization
Varythe wingloadingbyplus/minus20% andthe aspectratioby plus/minus20% to
determine the optimumcombinationusingthe carpetplotmethodof Chap.19.
3.9. Sizes and Capacities
3.9.1. Fuselage
Fuselage Length:25.86 [𝑓𝑡]
Fuselage maximumdiameter:8.62[𝑓𝑡]
3.9.2. Wing
Wingspan:37.34 [𝑓𝑡] Root chord:7.66 [𝑓𝑡]
Surface area: 178.76 [𝑓𝑡2] Root chordthicknessratio:15%
Wettedarea:732.92 [𝑓𝑡2] Tip chord:1.91 [𝑓𝑡]
Taper ratio:0.25 Tip chordthicknessratio:12%
39. LE Sweepangle:2[degree] MAC length:5.36 [𝑓𝑡]
Aspectratio:7.8 MAC location:7.47 [𝑓𝑡]
3.9.3. Tail
- Horizontal Tail
Root chord:5.41[𝑓𝑡] Aspectratio:5.2
Tip chord:1.35 [𝑓𝑡] Arm length:12.93 [𝑓𝑡]
Span:17.60 [𝑓𝑡] Taper Ratio:2.5
Area:59.55 [𝑓𝑡2]
- Vertical Tail
Root chord:8.39 [𝑓𝑡] Area:41.29 [𝑓𝑡2]
Tip chord:2.10 [𝑓𝑡] Aspectratio:1.5
Span:7.87 [𝑓𝑡] Arm length:12.93 [𝑓𝑡]
Taper ratio0.25
3.9.4. Landing Gear
The landinggearare designedtohave atotal addedheightof 16 inchesto the aircraft.
Witha tricycle type gearconfiguration,eachstrutwill have asingle Type III(low pressure)
wheel. More detail willbe giveninalaterdiscussion.
3.9.5. Fuel
The fuel systemiscapable of holding986 lbf of Jet-A fuel or147 gallons. More detail will
be givenina laterdiscussion.
3.9.6. Power Plant
The FlashfeaturestwoDGEN 380 enginesmountedaftof the wings. Each engine weighs
175 lbf andis 4 feet,5 inchesinlength. More detail will be giveninalaterdiscussion.
40. 3.10. Weight and Balance
The weightandbalance of an aircraftis of upmost importance. Itisimportantfromthe
verybeginningof the flightuntilthe aircraftisbackon the ground. Properweightandbalance
ensuresthe safetyof the flightandallowsease of maneuverability. The operatorof a light
aircraft suchas the Flash will needtocloselymonitorthe weightandbalance throughoutthe
flight’sentiretyasthe limitscanbe easilyexceededandthushave detrimental effects. The
followingderivationsare baseduponstatisticsandchapter15 of the Raymer text. Note:For the
mostaccurate information,the aircraftmustbe builtandweighedforaproperweightand
balance to be derived.
Withthe diagramabove,the followingtable wasformulatedasastatistical model of the
expectedweightandbalance of the Flash.
41. Weight
lbs
Loc
ft
Moment
ft-lbs
Weight
lbs
Loc
ft
Moment
ft-lbs
Structures 2296.9 35037.5 Equipment 466.84 7460.72
Wing 1109 14.5 16080.5 Flightcontrols 105 15 1575
Horizontal tail 130.5 30 3915 Hydraulics 15 0
Vertical tail 78 30.5 2379 Pneumatics 7 11 77
Fuselage 787 13 10231 Electrical 180 23 4140
Main landing
gear 130 16 2080 Avionics 45 4 180
Nose landing
gear 40 6 240 Furnishings 80 12 960
Firewall 22.4 5 112
Air
conditioning 39.84 8 318.72
Emptyweight
allowance 10 21 210
Propulsion 566 13135
Total weight
empty 3329.74 16.69 55633.22
Engines -
installed 448 25.5 11424
Fuel
system/tanks 118 14.5 1711 Useful load 1568 20918
Crew 150 9 1350
Fuel - usable 946 14.5 13717
Fuel - trapped 10 14.5 145
Oil 12 25.5 306
Passengers 450 12 5400
Takeoff gross
weight 4897.74 15.63 76551.22
3.11. Performance and Sub-System Designs
The designof each of the subsystemsare inaccordance with§23 of Federal Aviation
RegulationsforAviationMaintenanceTechnicians(FARAMT). Thisisonlya general overviewof
the equipmentandsystemoperationof the majorsubsystemsandisnotall inclusive. Thatisto
say thissectiondoesnotoutline all of the requirementslaidforthinthe FARAMT.
3.11.1. Flight Controls
Flightcontrolsare essential tocontrol the aircraftinall aspectsof flight. The flight
controlsmodifythe aerodynamicsurface of the wingandin turnchange the liftanddrag
42. producedby the surface it affects. The resultrotatesthe aircraftaround one of three,or a
combinationof the three,axestochange the flightpathof the aircraft. The three axesand the
correspondingflightcontrolsare the lateral axis,longitudinalaxisandvertical axis
correspondingtothe pitch,roll andyaw controlsrespectively. Pitchcontrol utilizesthe
horizontal stabilizer(horizontal tail surface),roll control utilizesthe ailerons(controlsurface
hingedonthe trailingedge of the wings),andyaw control utilizesthe vertical stabilizer(vertical
control surface attachedto the trailingedge of the vertical tail).
The Flashwill feature the mostcurrentandpilotfriendlycontrol surfacesthatwill create
the ease and comfortof flight. The Flash will be usingdifferentialpressureailerons,where one
ailerongoesupthanthe otheraileronwill deflectdown. Thiswillcreate amore significant
change in liftanddrag and a strongerroll overthe longitudinalaxis. The aileronswillalsobe
slotted,inordertoadd additional energytothe boundarylayer. Atthe trailingedge of the
aileronswill be trimtabs,alsocontrolledbythe commandof the pilotinthe cockpit. These are
small movable portionsof the control surface thatalterthe camber of the wingsothat the
change in the deflectionwill holdthe aircraftinanaerodynamicforce. There will be balance
tabs locatedonthe same control surface asthe trimtabs; the ailerons. Thistabaidsin the
movementof thissurface. The flapswill be afowlerflap. The fowlerflapisatype of slotted
flap. Thisflapwill change the camberof the wingandalsoincreasesthe wingareaby slidingthe
flapbackwardson tracks. The Flashwill use afullymovable horizontal stabilizerwithanti-servo
tabs. The anti-servotabisinstalledonthe trailingedgeof the control surface andassistsin
holdingthe control surface initsnew positionratherthanhelpingitmove. Thiswill decrease
the needforadditional actuators. There will be aconventionalhingedrudderlocatedonthe
trailingedge of the vertical stabilizer. These control surfacesare operatedthroughphysical
commandsfromthe cockpitcontrolsmade bythe pilot. These commandsare relayedtothe
43. flightcontrol surface throughseveral differentpossible meansincludingmechanical,hydraulic
and fly-by-wire.
The Flashwill be usinga fly-by-wire system. Fly-by-wire,intermsof ourapplicationis
an electrical primaryflightcontrol system(EPFCS)whichisdefinedbythe UnitedStatesAir
Force as, “a flightcontrol systemmechanizationwhereinthe pilot’scontrol commandsare
transmittedtothe momentor force produceronlyviaelectrical
wires.” The keyfeaturesthatare associatedwithfly-by-wire systemsare the replacementof
heavyhydraulicsystemswithelectrical wiresandcomputerassistedautostabilizers. The fly-by-
wire systemreducesthe fuel costs,increasepassengercapacity,haslowermaintenance costs,
44. improvesflightefficiencyandreducesthe fatigue of the pilot. All flightandtrimcontrolsgo
througha transducer,where itwill roll orpitch,andphysical commandsbecome encoded. The
encodedinformation issenttothe control computerwhichdeciphersthe informationandsends
out commandsto the surface actuators. The control computeralsocontainsaircraftmotion
sensors,whichisalsotakenintoaccountand makesadjustmentssothe pilotdoesnothave to
conduct extraworkto fulfillthe flightpathhe wants. The commandoutputfrom the control
computeralso goesthroughservovalvesattachedtothe actuator. The EPFCSfly-by-wire
systemcan containmultiplelayersof redundancytoincrease itsreliability,withoutthe tradeoff
weight,costandmaintenance.
The fly-by-wire systemwill containbuiltintestequipment,whichwillquicklydetectand
45. isolate failuresinthe system. Thisplacesanaddedlayerof safetyandalsodecreasesthe
amountof maintenance manhoursbydirectingthe mechanictothe source of the failure. North
AmericanRockwell Corporationestimatesthata fly-by-wire systemwilldecreasethe downtime
of an aircraftby at least3% and a reductionincontrol systemandmaintenance manhourscan
be reducedbyas much as 80% or more. With a fly-by-wiresystem,the control computercan
improve aircrafthandlingqualitiesbyadjustingthe stickfeel tothe pilotspreference forall flight
conditions. Thisisdue tothe control andstabilityaugmentation. Control augmentationis
referencedto the removal of the mechanical linkfromthe pilottothe seriesof servosforthe
fly-by-wireoperations. Pilotinputissenttothe commandmodel anddata fromthe aircraft
motionsensorsare thencompiledandsenttothe servoamplifier. The seriesservoreceives
informationfromboththe servoamplifierandthe frictionhysteresisbefore sendingthe
compiledinstructionstothe surface actuator. The stabilityaugmentationisoftenreferredto
the damper. Aircraftmotionsensorssenddatato a servoamplifierwhichisthensenttoa series
servo. The pilotinputissentto the frictionhysteresisandisalsosentto the seriesservobefore
all the informationissenttothe surface actuator. Both formsof augmentationare throughthe
same fly-by-wire system. Inthe fly-by-wiresystem,the seriesservoisprotectedbyavalve from
the aircraft motionsensorsandthe surface actuator isalsoprotectedbya valve so as to not
cause structural damage to the mechanism.
The primaryflightcontrol computeristhe PFCC-4100 fromRockWell Collins,Inc. Itwill
be locatedinthe nose of the aircraft. It offersveryhighintegritycommandoutputstothe
actuationsystem. Itwill have stable augmentationandenvelope protection. Itcoordinates
flightcontrol systemmaintenance toensure the qualityof the flightcontrol system. Thisflight
control computerismulti-channeledtoevaluate manyinputs,whichallowsone computerto
operate manyredundantsystems. Thiswillensure the safetyof the flightcontrol system. The
46. actuators will be fromMoog. The primaryflightcontrolswill be customizedfly-by-wire thatwill
come in dual redundantdesigns. All redundantsystemswill be ranthroughthe primaryflight
control computer. All informationthatgoesthroughthe flightcontrol computeroriginatesfrom
the commandsgeneratedinthe cockpitandfrom the cockpitcontrol. These controlsinclude
the control column,side stickandpedals.
The designof a fly-by-wire cockpitlayoutisdeterminedonthe intendeduse of the
aircraft. Dependingonpurpose the customermaychoose the control columnor side stickas
part of theiroptionsandthe panel designdependsontheirneeds. The control columnis
suggestedfortrainingpurposeswhilethe side stickisrecommendedforexperiencedpilots. The
control columnstyle isrecommendedforthe trainerdesign. There are control actuatorsthat
provide realisticfeedbacktothe pilotsotheymayexperiencemaneuvers. Thisconfiguration
doesincrease weightandrequire more roomthanthe side stickconfiguration. However, for
trainerpurposes itisrecommendedtoteachnew pilotsfeedbackfromtraditional control
columns. The side stickstyle isrecommendedforexperiencedpilotsforitsease of use and
reducedweight. The side stickisadaptedforemergencysituationsandpreventsthe pilotfrom
performingmaneuversoutside of the aircraft’scapabilities. Due toits positionandsmall size,
the side stickismore comfortable andprovidesanunobstructedviewof the control panel.
47. Cockpitpanelsare arrangeddependingonuse andneedof the pilot. The locationof the
controlstakesintoaccount eachsystems’importance,the frequencyof asystem’soperation,
the ease that the controlscan be reachedand the shape of the control.
3.11.2. Avionics
The Avionicspackage inThe Flashisprimarilysuppliedbythe GarminG1000. The G1000 is
the premiere glasscockpitandthe industryleaderincrew resource managementandreliable
operation.Whenselectingthe avionicspackage,Garminpresentsthe mostrobustoutof box
solution,whichincludesbuiltinredundancies,anefficientuserinterface,andthe modular
abilitytoinclude awide varietyof auxiliaryunits,calledLine ReplaceableUnits (LRUs).In
additiontothese builtin redundancies,ouravionicspackage will alsoincludeathirdlayerof
redundancy,ina small trioof traditional mechanical gaugesoperatingona completelyseparate
subsystem.
Image: GDU 1040
The G1000 displaysall of itsinformationthroughitstwo10.4 inchdisplays,one of which
isa PrimaryFlightDisplay (PFD),andthe otherisdesignatedthe Multi-FunctionDisplay (MFD).
Theyare boththe same unit,GDU 1040, and theyare designatedbasedupontheirphysical
locationinthe cockpit.The PFD isthe unitdirectly infrontof the pilotposition,andthe MFDis
48. infront of the copilotposition.Theyare bothfullycustomizableastowhat can be displayedon
eitherscreen,andare redundanttoeach other.Inthe eventof failure all pertinentinformation
be displayedonanysingle screen.Thesedisplayspresentinformationsuchasthe artificial
horizon,heading,VORheading,windspeed,andengine outputs,amongotherinformation.
Image: GRS 77 AHRS
The firstlevel of informationprocessingisthe GRS77 Atitude,Heading,andReference
Unit (AHRS).Itreceivesinputfromthe GMU 44 Magnetometer,aswell asit'sbuiltintiltsensors,
accelerometers,andrate sensors.The AHRSisthe primarysource foraircraft attitude andflight
characteristicsinformation.
49. Image: GDC 740 ADC
The primarycomputingcenterof the systemisthe GDC 740 AirData
Computer(ADC).The ADCreceivesthe inputfromthe Pitot-Staticprobe,the GTP59 OAT Probe,
as well asthe AHRS and IntegratedAvionicsUnits.The ADCdeterminessevenprimary
parameters:Total AirTemperature,PressureAltitude,IndicatedAirspeed,CalibratedAirspeed,
Vertical Speed(Rate of Climb),andMach.
The GEA 71 Engine/Airframe Unit(EAU) providesthe systemwithconnectiontothe
enginesFADECandairframe sensors,suchasfire sensors.Itcommunicateswiththe systemby
RS-485 digital communicationlines.
50. Image: GIA 630 IAU
The heart of the systemare the twoGIA 630 IntegratedAvionicsUnits(IAUs).The IAUs
provide the displayswiththeirfunctionality,aswell ascontainthe GPS receiver,NAVradio
receiver,andcommunicationstransceiver.The twounitsprovideeachotherwithredundancy.If
one unitfails,the othersensesthisfailure,andall tasksare handledbythe functioningunituntil
the failedunitisreplaced.Thisredundancycoversall functionswiththe exceptionof GPS,which
requiresboth unitstoachieve the requiredaccuracy.
The IAUs communicate withthe othercomponentsthroughavarietyof
communicationlines.The displayscommunicate toeachother,aswell asthe IAUsthrough
standardEthernet.The IAUs communicate toall of the othercomponentsthroughARINC429, as
well asRS-232. The use of twocommunicationstylesprovidesautomaticerrorcorrection
throughcomparison.
The primaryaudiointerface forthe G1000 is the GMA 13470 AudioPanel.Itcontrols
all audiocontrols,includingintercomradios,NAV radio,communicationsradios,andoptional
XMradio.It is mountedbetweenthe displays,andcommunicatesonlywiththe IAUsacrossRS-
232.
51. The GTX 330 Transponderisa Mode-StransponderwhichprovidesmodesA,C,and S
ATC communication.Itiscontrolledbythe IAUsthroughthe display.
There are manyoptional LRUs whichcan be incorporatedintothe G1000, whichcan add
a varietyof featuresandprovide more robustfunctionalitybaseduponthe customer'sneeds.
Additional features includeXMRadio,Weathersystems,andanynumberof differentdisplays.
The basic configurationof the G1000 providesthe necessaryfunctionalitytofullyequipan
aircraft forflight,andrepresentsthe cuttingedge of modernglasscockpit.
Images: Artificial Horizon, Indicated Air Speed, Pressure Altitude
In additiontothe G1000, the aircraft will alsobe equippedwithacompletelyseparate set
of traditional units.These unitsinclude Artificial Horizon,IndicatedAirSpeed,andAltitude
displays.Theyare includedinouravionicspackage tocreate a thirdlayerof redundancy,which
will protectthe aircraftinthe eventof a full G1000 systemfailure.
3.11.3. Electrical System
The electrical systemonThe Flashisa parallel-type bussystemoperatingat400 Hz 115
VoltsAlternatingCurrent,inone of three phases,and28 VoltsDirectCurrent.The parallel bus
arrangementallowsforimmediate failuredetectionandpreventsthe aircraftfromlosingfull
52. powerinthe eventof incidentssuchassingle enginefailure.Itconsistsof twoparallel
subsystems, namedLeftandRight,whichisnamedbaseduponthe enginepoweringthe
subsystem,asviewedfromthe pilot'sperspective.The LeftandRightsystemsare connectedat
twopoints,the firstisthe Main AC Bus,and the secondis the Main DC Bus.The system also
allowsforthe inputof a groundpowercart or truck, whichisa separate ACsubsystem.
In the eventof an engine orcomponentfailureinanyone ACsubsystem, the
functioningsubsystemcanautomaticallyandquicklytransferitspowerintoasecondary
subsystem.The secondarysubsystemscontainmanyof the same componentsasthe main
subsystem,butfeedcompletelyseparate buses.These buses,whileseparate,canpowerall of
the essential componentstothe aircraft.
The engine'sprimaryelectrical outputcontrol isit'sincorporatedFull Authority
Digital Engine Control (FADEC) software.The firstcomponentwhichphysicallybeginsgenerating
electricityisthe IntegratedDriveGenerator (IDG),mountedontothe engine,andcontrolledby
the FADEC. The engine produceselectricitybasedonitscurrentoperatingconditions,andthe
it's the responsibilityof the FADECtocontrol the IDG inorder to preventpotentiallyharmful
situationsfromconcurring.The IDG utilizesthe GeneratorControl CurrentTransformer(GCCT),
communicatingthroughaloadcontroller,toconditionthe ACpowerintothe acceptable range.
The GCCT transformsthe AlternatingCurrent (AC)powerdirectlyfromthe generator
into400 Hz 115 VoltsAC (VAC),andintothe appropriate phase withthe restof the system.The
initial phase isdeterminedbythe firstpowersource operatingduringthatruncycle,andall
otherAC systemsconformtoitfor the durationof the run cycle.
The GeneratorBreaker(GB) isthe firstbreakerinline from the generator.Itsprimary
functionisto preventcommonfailuresfromaffectinganycomponentswhichare still operating.
53. The GB alsoprovidesthe connectionbetweenthe MainACBus,the IDG, andthe secondaryAC
bus.
The Bus Tie Breaker(BTB) provides the cutoff pointforthe mainACsubsystem, and
closesthe circuitincase of failure.The BTB,GB, andIDG all communicate withthe GCU,which
monitorsthe currentflowinthe subsystemandcan determinefailure.
The DifferentialProtectionCurrentTransformer(DPCT) isacomparative transformer
whichusesthe methodof differentialprotectiontomonitorthe currentflowingthroughboth
the primaryand secondarysubsystems,aswell asthe IDG output.While the GCCTmonitorsthe
systemasa whole,andcommunicateswiththe LoadControllerandGCU, the DPCT monitorsthe
miscellaneousfeederwiresforshortedandopenconditions.The purpose of the DPCTisnot
necessarilytotransformthe engine output,butitregulatesthe componentsdrawingfromthe
system.
Completelyseparate fromeitherof the IDGpoweredlinesisthe GroundPower
subsystem.Thisline allowsforthe use of acart or truck to supplythe aircraftwithpower,when
it isavailable.There are manyadvantagesforutilizingthe groundpoweravailability,primarily
the abilitytostart the aircraft's engineswithoutthe needtodraw energyfromthe battery.The
subsystemcontainstwoDPCTsforline protection,andanEngine Pressure Ratio(EPR) sensor,
whichisprovidedthe currentengine output.ThisEPRsensorclosesthe subsystemwhenthe
engineshave reachedaself-sustainingoperation,whichprotectsthe groundcart fromdamage
due to substantial backloads.
The Direct Current(DC) subsystemisagaindividedintothree furthersubsystems,
baseduponwhere theirpowerisrectifiedfrom.The primaryDCsubsystembeginsatthe
Transformer/RectifierUnit(T/RU) main,whichdrawsdirectlyfromthe MainAC Bus.Similarly,
54. the T/R U Leftand Rightdraw theirpowerfromthe correspondingACBuses.All power
conditioningforthe DCsubsystemsishandledbythe T/RUs, whichconvertthe ACpowerinto
24 V DC power.
Surge protectioninthe DC subsystemsisprimarilyprovidedbysemiconductor
diodes.These diodespreventDCpowerfromtravelinginareverse pathof the intendedflow,as
well asprotectthe systemfromexcessivevoltages.Fuses
From the T/R U powerisroutedintothe Essential DCBus,whichpowerscomponents
such as the BatteryBus, andthe avionics.It'splacementbefore the MainDCBus createsa
possibilityforprotectionfromfailure whichmayoccur inthe Main DC Bus. Thisarrangementis
an attemptto limitthe effectof,say,lightfailure,fromaffectingflightessential components
such as the avionics.
The Battery Busprovidespowerfromthe batteryintothe Essential DCBuswhen
there isneedforit.The mostcommonneedforbatterypoweriswhenthere isno groundcart in
place for the systembefore andduringenginestart.Batterypowerisalsoa final backupfor
flightinthe eventof dual engine failure,andaccordingto FAA regulationsthe batterymustbe
able to provide powerforthe aircraftfor30 minutes.The aircraftwill containtwo12 V Lithium-
Ionbatteriesconnectedinseriestofulfill thisrequirement.
While Lithium-Ion(Li-Ion)batteriesare still relativelynew inaviation,theiruse has
become more accepted.CompaniessuchasEaglePicherhave fullydevelopedFAA registeredLi-
Ionbatteriesandchargers,whichprovide asubstantial increase overtraditionalLead-Acid
batteriesinthe areasof weight,cycle variability,anddischarge duration.
55. The batterybus alsohousesthe batterycharger,whichreceives powerfromthe AC
subsystems.While physicallyseparate fromthe DCsubsystem, the batterychargerconverts AC
powerintoDC powerinorder to charge the batteries,whichthenpowersthe DCsystemas
describedinthe conditionsabove.Whilethe batteriesstore 24V of power,the batterycharger
provides28 V of powerto charge the batteries.
56. Image: Electrical System Diagram
3.11.4. Landing Gear
The Flashutilizesatricycle type landinggearsystemwithasingle wheelperstrut. This
type of gear will allowthe cockpittoremainata level attitudeduringtaxi andtakeoff aswell as
57. allowthe pilotgoodvisibilityandcontrollability. The landinggeariscapable of withstandingup
to 90% of its max takeoff grossweightinthe eventof anemergencylandingneedingtobe
performedshortlyaftertakeoff.The landinggearwill be retractable toreduce the effectsof
drag and allowasmoother,fasterflight. The landinggearutilizesa12 VDC bi-directional
electro-hydraulicpowerpackandpumpto place the gear intothe desiredposition. The landing
gear will be producedandassembledbyHeroux-DevtekIncorporationandshippedtousfor
final installationontothe aircraft.
The wheel andtire selectionsare basedupontables,chartsandequationslistedin
chapter11 of the Raymertext. The landinggearwill weigh130 lbf.The nose gear will have a7 ̊
forwarddisplacementtocounteract any tendencyforthe gearto retract upona hard landing.
Both the nose andmain landinggearwill utilizethe same size tiresandwheels. The tire will be
type III,lowpressure,andcan supporta maximumspeedif 120 mph,or 104 knots. The tiresare
capable of supporting4400 lbf,90% of the takeoff weight. The areafootprintof the tire is90
in². The tireswill have anoverall diameterof 25.65 inchesand a widthof 8.7 inches. The tires
will be capable of holdinga maximumof 55 psi. During landings,acenteringcamwill ensurethe
nose gearis inthe straightaheadpositionsince itisthe onlygearcapable of swiveling.
Additionally,adrag strut andside brace linkwill be utilizedonthe maingearforsafetyconcerns
inthe eventof a highcrab landing. Air-oleotype shockabsorbersare utilizedoneachstrutand
can be servicedviaanairvalve at the top of each strut. The use of an air-oleoversusaspring-
oleoallowsforconservationof weightwhile still cushioninglandingsandtaxiingoverrough
surfaces.
Whenthe pilotselectsthe landinggeartomove intoeitherthe extendedorposition,an
electricmotorisenergizedandrotatesacam plate that opensthe landinggearstowage doors,
58. positionsthe gear,andclosesthe doors. Once the gear isin the selectedposition,amicroswitch
breaksthe circuitto the motorand causesthe appropriate gearindicationtobe displayedon
the multi-functionaldisplays. The gearwill retractto stow in a fuselage-poddedconfiguration.
For the purpose of enhancedsafety,alanding-gear-positionindicatorsystemisutilized. Squat
switchesallowthe systemtodeterminewhenthe aircraftisonthe ground,disallowingthe gear
to be retractedaccidently. A warninghornwill soundwhenthe throttle isreducedbelow 100
knotsand the landinggearisnot inthe downposition. Inthe eventof acomplete electrical
failure,abackupCO2 accumulatorwill use itscharge to place the landinggearin the landing
configuration,referredtoasan emergencygearuprelease valve blow downsystem.
The brakesare a single disktype andare operatedviaa brake-by-wire system. When
the pilotpressesthe brakes,anelectrical signalissentfromthe brake pedal transducersandthe
Garmin 1000 systemtoactuate electrical brake actuators. Thissystemutilizesnohydraulic
fluid,allowingforweightconservation. The brake actuatorsprovide brakingpowertoeither
one or all wheels,atthe pilot’sdiscretion,viapressure appliedtothe individual footpedals.
Disksare rigidlyboltedtothe wheel andabrake housingisattached. The pistonsinthe brake
housinghave liningsonthemwhichmustbe replacedwhenwornbelow tolerances,muchlike
the brakesof a car.
3.11.5. Pressurization System
Environmental systemistypicallyincludingairconditioningsystemandpressurization
system.Theyare workingtogethertocreate a comfortable atmosphere forpassengersand
crewsin the cabin.
For personal lightjetand/orverylightjet,cabinpressure differential isgenerallyupto
6.7±0.1 psi.The presetpressure differential valueis6.8psi.Thisallowsasealevel cabin
59. altitude upto12,000 feet.Andourmaximumcruise altitude is25,000 feet,sothe cabin
altitude wouldbe 5,000 feet.
The basic componentsincludeanavionicslinkeddigital controllerandtwooutflow
valvesmountedinthe aftpressure bulkhead.The MFDdisplaysall pressurization
parametersandthe PFDs provide pilotinterface forentryof landingfieldelevation.Inthis
design,nobleedingairsystemisappliedinsteadof conventional bleedingairsystem
coordinatedwithpneumaticsystem.Firstly,cabinairwill be venteddirectlyfromthe
outside throughdedicatedinletsoneachside of the plane'sbellyandwill notpassthrough
the engines.Andthenelectricallydrivencompressorcompressesthe ramlow-densityair.
Afterthat itis transportedviaductsto the air conditioningpacks.Withinthe A/Cunit,the
desiredtemperature isachievedbyregulatingthe adjustable speedmotorcompressorsat
the requiredpressure withoutsignificantenergywaste.Andthe regulatedairdistributes
throughoutletsinthe cockpitand overheadventsinthe cabin,respectively.The system
60. may be operatedanytime inflight,oronthe ground whengroundpowerisconnectedor
eitherengine isrunning.A freshairventwitha bloweranda checkvalve islocatedbeneath
the nose baggage compartmentto provide outsideairtothe cockpitwheneverthe cabinis
not pressurized.
Thisapproach issignificantlymore efficientthanthe traditional bleedsystembecause it
avoidsexcessive energyextractionfromengineswiththe associatedenergywaste by
pre-coolersandmodulatingvalves.Thatresultsinsignificantimprovementsinenginefuel
consumption.
3.11.6. Fire Protection System
Fire isone of the most dangerousthreatstoan aircraft.Fire protectionsystemsisvery
importantforeveryaircraft.It isinstalledinanaircraftto detectandprotect againstan
outbreakof fire.Forthe fire zonesfor our aircraft “The Flash”it will be dividedintothree
sections.Thisinclude the engine section,the nose compartmentandthe maincabin.Forthe
61. engine section,itwill be dividedintotwozonesnamelyzone A andzone B.Zone A isgoingto
coverthe core sectionof the engine anditis alsogoingto be providedwithfire detectionand
extinguishing.Zone Bwill coverthe exhaustpipeandpylonsection.One extinguisherisgoingto
be paced oneach engine withHalon3301 andone isgoingto be placedinthe cockpit.
For the air cooledradial engines,the powersectionandall portionsof the exhaustsystemmust
be isolatedfromthe engine accessorycompartmentbyadiaphragmthat meetsthe firewall
requirementsof part23.1191. The designof the fire protectionforthisaircraftwill be in
compliance withthe requirementwhichinclude:
(a) Each engine,powerunitsandall othercombustionequipmentwillbe
isolatedfromthe aircraftbyfirewalls.
(b) The firewall willbe constructedsothatno hazardousquantityof liquids,
gas, or flame canpass fromthe compartmentcreatedbythe firewall.
(c) Each openinginthe firewall will be sealedwithclosefitting,fireproof
grommets,bushing,orfirewall fittings.Ourfirewall will be made upof composite
material.The firewall forthisaircraftwill be protectedagainstcorrosionandalsowill
be a fireproof andthisisgoingtoprotectit fromany dangerof fire andas a result,
passengersandcrewdoesn’tgetelectrocutedwhentheyinsideof the cabin.For
example the material thatwillbe usedwhichrequiresfirewall materialsandfittings
mustresistflame penetrationforatleast15 minutes.Forthe designforthisaircraft,
the followingmaterial will be used
1. Stainlesssteel sheet,0.015 inchthick
2. Mildsteel sheet(coatedwithaluminumorotherwise
protectedagainstcorrosion) 0.018 inch thick
62. 3. Monel metal,0.018 inchthick
4. Steel orcooperbase alloyfirewall fittings
5. Titaniumsheet,0.016 inchthick
All aircrafthave an extinguishingsystem.The kindof extinguisherthatisgoingtobe usedonthis
aircraft will be the classB whichismore effectivewithflammableliquidsandwithchemicals
that include monoammoniumphosphateandsodiumbicarbonate.The nexttype of extinguisher
that will be usedisclassC whichissuitable forfire inelectrical equipmentwithchemicalsthat
include monoammoniumphosphate andsodiumbicarbonate.
Image: Fire Suppression Bottles for Engines
63. 3.11.7. Fuel System
All powered aircraft require fuel on board to operate the engines throughout the phases of
flight. A fuel system consists of storage tanks, pumps, valves, filters, fuel lines, monitoring
devices, and metering devices. Each system must provide an uninterrupted flow of contaminant
free fuel regardless of the aircraft’s attitude or flight condition. Varying fuel loads and shifts in
weight during maneuvers must not negatively affect control of the aircraft in flight. In general,
fuel systems must be constructed and arranged to ensure fuel flow at a rate and pressure
established for proper engine functioning under each likely operating condition. It also must be
designed and arranged to prevent the ignition of fuel vapor within the system by direct lightning
strikes.
For multiengine aircraft, each fuel system must be arranged so that, in at least one system
configuration, the failure of any one component does not result in the loss of power of more than
one engine. If two fuel tanks interconnected to function as a single fuel tank, there must be
independent tank outlets for each engine, and each incorporating a shut-off valve. The shutoff
valves may serve as firewall shutoff valves. Lines from each tank outlet to each engine must be
completely independent of each other. The fuel tank must have at least two vents arranged to
minimize the probability of both vents becoming obstructed simultaneously. In addition, aircraft
fuel tanks must be designed to retain fuel in the event of a gear-up landing. In case of sever
emergency situations, there must be a means to allow flight crew members to rapidly shut off the
fuel to each engine individually in flight.
64. The Flash has two fuel tanks that can carry a combined 147 U.S Gallons of Jet A fuel,
consisting of one tank per wing. Each wing has fuel receptacle that is located above the wing root
behind a spring loaded cover flap. Each receptacle then consists of a fueling nozzle adapter and
sealing cap. From the receptacle a fuel line runs downward into each respective fuel tank. There
are two primary fuel pumps in each tank located at opposite sides of the respective tank to allow
for continuous supply of fuel to the engine during maneuvers when the aircraft’s attitude is not
level. These two fuel pumps flow into one singular fuel line at a T-joint with one-way valves
preventing backflow returning to the fuel tank. Secondary or backup fuel pumps are located
adjacent to the primary fuel pumps; one secondary pump for each primary fuel pump. They use
65. most of the same fuel lines as their adjacent primary pump. The secondary pumps are on standby
until activated by the pilot, or if fuel pressure drops below a certain amount, they will be
automatically switched on. A collector box in the wing root keeps the electrical pumps inlets
submerged. To prevent pump cavitation, a pump and flaps valves ensure enough fuel in the
collector box at all times.
A single fuel line connects each tank with a crossfeed valve located along the centerline of
the fuselage. An air valve located above the fuel pump allows air to be vented outside for priming
the crossfeed line at engine startup, and allows for air to be pumped into the crossfeed line at
engine shutdown to prevent unwanted expansion of fuel during times of engine inactivity. There
is also a fuel vent system with vent tanks located at the wing tips which prevent damage to the
wings due to excessive buildup of positive or negative pressures inside the fuel tanks and to
66. provide ram air pressure within the tanks. For fuel indication within the cockpit, four fuel sensors
are installed inside each tank, and are equally spaced across the full length of the tank. measure
fuel levels at each sensor’s location and send the information to a computer that constantly
calculates the overall fuel level of the tank. For manual measurement, there are direct measuring
sticks located on the wings.
4. MANUFACTURINGPLAN
4.1 Manufacturing Readiness Levels
Matters of manufacturingreadinessandproducibilityare asimportanttothe successful
developmentof a systemasthose of readinessandcapabilitiesof the technologiesintendedfor
the system.Theirimportance haslongbeenrecognizedinthe Departmentof Defense(DoD)
acquisition,andare reflectedincurrentDoDacquisitionpolicies.Foranaerospace company,itis
verybeneficial tofollow the DoDstandardsandpractices.
4.1.1 Defining Manufacturing Readiness
Accordingto the DoD, ManufacturingReadinessisthe abilitytoharnessthe
manufacturing,production,qualityassurance,andindustrial functionstoachieve anoperational
capabilitythatsatisfiesmissionneedsinthe quantityandqualityneededbythe aircraftto
performas itis designedtoatthe "bestvalue."Bestvalue referstoincreasedperformanceas
well asreducedcostfor developing,producing,acquiring,andoperatingsystemsthroughout
theirlife cycle. Timelinessalsoisimportant.Ouraircraft,"The Flash"mustmaintaina
technological advantageoverourcompetitor'saircraft.Thisrequiresefficientdevelopmentand
acquisitioncyclesforadvancingtechnologies.
ManufacturingReadinessbeginsbefore,andcontinuesduringthe developmentof an
aircraft's systems,andcontinuesevenafterasystemhasbeeninthe fieldforanumberof years.
67. The abilitytotransitiontechnologysmoothlyandefficientlyfromdevelopment,production,and
deploymentintothe fieldisacritical enablerforevolutionaryacquisition.
ManufacturingReadinessLevels(MRLs) are designedtobe measuresusedtoassessthe
maturityof a giventechnologyfromamanufacturingprospective.The purpose of MRLsare to
provide decisionmakerswithacommonunderstandingof the relative maturity,andattendant
risksassociatedwithmanufacturingtechnologies,products,andprocessesbeingconsideredto
meetDoD requirements.
4.1.2 Manufacturing Readiness Levels
There are tenMRLs that are correlatedtonine TechnologyReadinessLevels(TRLs) inuse.
The ten MRLs are describedindetail below.Inregardstoproductionof the aircraft,at MRL 8,
lowrate initial productioncanbegin.AtMRL 9, there isthe capabilitytogo intofull rate
production.ByMRL 10, full rate productionisdemonstratedandleanpracticesforefficient
productionare inplace.
Accordingto the National AeronauticsandSpace Administration(NASA),TRLsare a type
of measurementsystemused toassessthe maturitylevelof aparticulartechnology.Each
technologyprojectisevaluatedagainstthe parametersforeachtechnologylevel andisthen
assignedaTRL ratingbasedon the projectsprogress.There are a total of nine technology
readiness levels.TRL1 is the lowestandTRL 9 isthe highest.
MRL 1: Basic Manufacturing Implications Identified
Thisis the lowestlevel of manufacturingreadiness.The focusistoaddress
manufacturingshortfallsandopportunitiesneededtoachieve programobjectives.Basic
researchbeginsinthe formof studies.
MRL 2: Manufacturing Concepts Identified
68. Thislevel ischaracterizedbydescribingthe applicationof new manufacturingconcepts.
Appliedresearchtranslatesbasicresearchintosolutionsforbroadly definedneeds.Typicallythis
level of readinessinthe Science andTechnology(S&T) environmentincludesidentification,
paperstudiesandanalysisof material andprocessapproaches.Anunderstandingof
manufacturingfeasibilityandriskisemerging.
MRL 3: Manufacturing Proof of Concept Developed
Thislevel beginsthe validationof the manufacturingconceptsthroughanalytical or
laboratoryexperiments.Thislevel of readinessistypical of technologiesincategoriesof
research,development,andmaterialsprocesseshave beencharacterizedformanufacturability
and availability,butfurtherevaluationanddemonstrationisrequired.Experimental hardware
modelshave beendevelopedinalaboratoryenvironmentthatmaypossesslimited
functionality.
MRL 4: Capability to produce the technology in a laboratory environment
In thislevel,technologiesshouldhave maturedtoatleastTRL 4. Thislevel indicatesthat
the technologiesare readyforthe developmentphase of acquisition.Atthispoint,required
investments,suchasmanufacturingtechnologydevelopment,have beenidentified.Processes
to ensure manufacturability,producibility,andqualityare inplace andare sufficienttoproduce
technologydemonstrators.Manufacturingriskshave beenidentifiedfor buildingprototypesand
mitigationplansare inplace.Targetcost objectiveshave beenestablishedandmanufacturing
cost drivershave beenidentified.Producibilityassessmentsof designconceptshave been
completed.Keydesignperformance parametershave beenidentifiedaswell asanyspecial
tooling,facilities,material handlingandskillsrequired.
MRL 5: Capability to produce prototype components in a production relevant environment
69. Thislevel of maturityistypical of the mid-pointinthe developmentphase of acquisition.
Technologiesshouldhave maturedtoatleastTRL 5. The industrial base hasbeenassessedto
identifypotential manufacturingsources.A manufacturingstrategyhasbeenrefinedand
integratedwiththe riskmanagementplan.Identificationof enablingcritical technologiesand
componentsiscomplete.Prototypematerials,toolingandtestequipment,aswell aspersonnel
skillshave beendemonstratedoncomponentsinaproductionrelevantenvironment,butmany
manufacturingprocessesandproceduresare still indevelopment.Manufacturingtechnology
developmenteffortshave beeninitiatedorare ongoing.Producibilityassessmentsof key
technologiesandcomponentsare ongoing.A costmodel hasbeenconstructedtoassess
projectedmanufacturingcost.
MRL 6: Capability to produce a prototype system or subsystem in a production relevant
environment
For MRL 6, technologiesshouldhave maturedtoatleastTRL 6. It isnormallyseenasthe
level of manufacturingreadinessthatdenotescompletionof S&Tdevelopmentandacceptance
intoa preliminarysystemdesign.Aninitialmanufacturingapproachhasbeendeveloped.The
majorityof manufacturingprocesseshave beendefinedandcharacterized,butthere are still
significantengineeringand/ordesignchangesinthe systemitself.However,preliminarydesign
of critical componentshasbeencompletedandproducibilityassessmentsof keytechnologies
are complete.Prototype materials,toolingandtestequipment,aswell aspersonnel skillshave
beendemonstratedonsystemsand/orsubsystemsinaproductionrelevantenvironment.A cost
analysishasbeenperformedtoassessprojectedmanufacturingcostversustargetcost
objectivesandthe programhasin place appropriate riskreductiontoachieve costrequirements
or establishanewbaseline.Thisanalysisshouldincludedesigntrades.Producibility
70. considerationshave shapedsystemdevelopmentplans.Long-leadandkeysupplychain
elementshave beenidentified.
MRL 7: Capability to produce systems, subsystems, or components in a production
representative environment
At thislevel,technologiesshouldbe ona pathto achieve TRL 7. Systemdetaileddesign
activityisunderway.Material specificationshave beenapprovedandmaterialsare available to
meetthe plannedpilotlinebuildschedule.Manufacturingprocessesandprocedureshave been
demonstratedinaproductionrepresentative environment.Detailedproducibilitytrade studies
and riskassessmentsare underway.The costmodel hasbeenupdatedwithdetaileddesigns,
rolledupto systemlevel,andtrackedagainstallocatedtargets.Unitcostreductioneffortshave
beenprioritizedandare underway.The supplychainandsupplierqualityassurancehave been
assessedandlong-leadprocurementplansare in place.Productiontoolingandtestequipment
designanddevelopmenthave beeninitiated.
MRL 8: Pilot line capability demonstrated; Ready to begin Low Rate Initial Production
Thislevel isenteringintoLow Rate Initial Production(LRIP)of the aircraft. Technologies
shouldhave maturedtoat leastTRL 7. Detailedsystemdesignisessentiallycomplete and
sufficientlystable toenterlowrate production.All materialsare availabletomeetthe planned
lowrate productionschedule.Manufacturingandquality processesandprocedureshave been
provenina pilotline environmentandare undercontrol and readyforlow rate production.
Knownproducibilityriskspose nosignificantchallengesforlow rate production.The engineering
cost model isdrivenbydetaileddesignandhasbeenvalidatedwithactual data.
MRL 9: Low rate production demonstrated; Capability in place to begin Full Rate Production
At thislevel,the system,componentoritemhasbeenpreviouslyproduced,isin
production,orhas successfully achievedlow rate initial production.Technologiesshouldhave
71. maturedto TRL 9. Thislevel of readinessisnormallyassociatedwithreadinessforentryintoFull
Rate Production(FRP).Allsystemsengineeringdesignrequirementsshouldhave beenmetsuch
that there are minimal systemchanges.Majorsystemdesignfeaturesare stable andhave been
provenintestand evaluation.Materialsare availabletomeetplannedrate production
schedules.Manufacturingprocesscapabilityinalow rate productionenvironmentisatan
appropriate qualityleveltomeetdesignkeycharacteristictolerances.Productionrisk
monitoringisongoing.LRIPcosttargetshave beenmet,andlearningcurveshave beenanalyzed
withactual data. The costmodel hasbeendevelopedforFRP environmentandreflectsthe
impactof continuousimprovement.
MRL 10: Full Rate Production demonstrated and lean production practices in place
Thisis the highestlevelof productionreadiness.Technologiesshouldhave maturedto
TRL 9. Engineeringdesignchangesare minimal,andgenerallylimitedtoqualityandcost
improvements.Systems,componentsoritemsare infull rate productionandmeetall
engineering,performance,qualityandreliabilityrequirements.Manufacturingprocess
capabilityisatthe appropriate qualitylevel.Allmaterials,tooling,inspectionandtest
equipment,facilitiesandmanpowerare inplace and have metfull rate production
requirements.Rate productionunitcostsmeetgoals,andfundingissufficientforproductionat
requiredrates.Leanpracticesare well establishedandcontinuousprocessimprovementsare
ongoing.
Althoughthe MRLs are numbered,the numbersthemselvesare unimportant.The
numbersrepresentanon-linearordinalscale thatidentifieswhatmaturityshouldbe asa
functionof where aprogram isin the acquisitionlife cycle.
72. Level Definition DoD MRL Description
1
Basic Manufacturing
Implications
Identified
Basic researchexpandsscientificprinciplesthatmayhave
manufacturingimplications.The focusis ona highlevel
assessmentof manufacturingopportunities.The researchis
unfettered.
2
Manufacturing
ConceptsIdentified
Thislevel ischaracterizedbydescribingthe applicationof new
manufacturingconcepts.Appliedresearchtranslatesbasic
research intosolutionsforbroadlydefinedmilitaryneeds.
3
ManufacturingProof
of Concept
Developed
Thislevel beginsthe validationof the manufacturingconcepts
throughanalytical orlaboratoryexperiments. Experimental
hardware modelshave beendevelopedin alaboratory
environmentthatmaypossesslimitedfunctionality.
4
Capabilityto
produce the
technologyina
laboratory
environment
Thislevel of readinessactsas an exitcriterionforthe MSA
Phase approachinga Milestone Decision.Technologiesshould
have maturedto at leastTRL 4. Thislevel indicatesthatthe
technologiesare readyforthe TechnologyDevelopmentPhase of
acquisition.Producibilityassessmentsof designconceptshave
beencompleted.Keydesignperformance parametershave been
identifiedaswellasanyspecial tooling,facilities,materialhandling
and skillsrequired.
5
Capabilityto
produce prototype
componentsina
production
relevant
environment
Mfg. strategyrefinedandintegratedwith RiskManagementPlan.
Identificationof enabling/critical technologiesandcomponentsis
complete.Prototypematerials,toolingandtestequipment,aswell
as personnel skillshave beendemonstratedoncomponentsina
productionrelevantenvironment,butmanymanufacturing
processesandproceduresare still indevelopment.
6
Capabilityto
produce a prototype
systemor subsystem
ina
productionrelevant
environment
ThisMRL is associatedwithreadinessfora Milestone Bdecisionto
initiate anacquisitionprogrambyenteringintothe EMD Phase of
acquisition.Technologiesshouldhave maturedtoat leastTRL 6.
The majorityof manufacturingprocesseshave beendefinedand
characterized,butthere are still significantengineeringand/or
designchangesinthe systemitself.
8
Pilotline capability
demonstrated;
Readyto beginLow
Rate Initial
Production
The system, componentoritemhasbeenpreviouslyproduced,is
inproduction,orhas successfullyachievedlow rate initial
production.Technologiesshouldhave maturedtoTRL 9. Thislevel
of readinessisnormallyassociatedwithreadinessforentry
intoFull Rate Production (FRP).All systemsengineering/design
requirementsshouldhave beenmetsuchthatthere are minimal
systemchanges.Majorsystemdesignfeaturesare stable andhave
beenprovenintestandevaluation.
73. 9
Low rate production
demonstrated;
Capabilityinplace to
begin
Full Rate Production
The system, componentoritemhasbeenpreviouslyproduced,is
inproduction,orhas successfullyachievedlow rate initial
production.Technologiesshouldhave maturedtoTRL9. Thislevel
of readinessisnormallyassociatedwithreadinessforentryinto
Full Rate Production(FRP).All systemsengineering/design
requirementsshouldhave beenmetsuchthatthere are minimal
systemchanges.
10
Full Rate Production
demonstratedand
leanproduction
practices
inplace
Technologiesshouldhave maturedtoTRL9. Thislevel of
manufacturingisnormallyassociatedwiththe Productionor
Sustainmentphasesof the acquisitionlife cycle.
Engineering/design changes are few andgenerallylimitedto
qualityandcost
improvements.System, componentsoritemsare infull rate
productionandmeetall engineering,performance,qualityand
reliabilityrequirements.Manufacturingprocesscapabilityisatthe
appropriate qualitylevel.
Accordingto NASA,the followingare the TechnologyReadinessLevelsmentionedabove are
displayedbelow.
74.
75. 4.2 Industrial Base
At thisstage,our ManufacturingReadinessLevel iscurrentlyatLevel 2,and then will be
proceedingintoLevel 3.
2
Manufacturing
ConceptsIdentified
Thislevel ischaracterizedbydescribingthe applicationof new
manufacturingconcepts.Appliedresearchtranslatesbasic
researchintosolutionsforbroadlydefinedmilitaryneeds.
3
ManufacturingProof of
ConceptDeveloped
Thislevel beginsthe validationof the manufacturingconcepts
throughanalytical orlaboratoryexperiments. Experimental
hardware modelshave beendevelopedinalaboratory
environmentthatmaypossesslimitedfunctionality.
Setupwill take approximately12-24 months.
Suppliers:
5 Price Induction(2DGEN 380 engines)
6 Garmin (Avionics)
7 Rockwell (FlightControls)
8 Héroux-Devtek(LandingGear)
4.2.1 Price Induction
Price inductionisone of the few companiestohave developedamodernaeronautical
gas turbine inthe pastdecade.Itsstate-of-the-artproductisthe DGEN 380 engine,the world’s
smallestturbofanintendedfor4-5 seatPersonal LightJets.Thishighbypassratiogeared
turbofanwasdesignedfromablanksheettoallow forthe adventof a new classof aircraftson
the general aviationmarket.Afterfifteenyearsof development,the engine isrecognizedasa
technical successandhas nowto enterthe certificationandindustrializationphase.
76. Price Induction’sadventure beganin1997, whenBernardEtcheparre,a French
entrepreneur,decidedtolaunchthe DGEN program to contribute tothe innovationinthe
general aviationmarket.Launchedasaventure project,withateamof youngengineers,the
program quicklygainedthe supportof Frenchaerospace laboratories,majorFrench
aeronautical companiesandinstitutional investmentfunds.
On October31st 2006, the firstDGEN 380 engine wassuccessfullyignitedwiththe test
benches.In2011, the DGEN 380 completeditsfirst150-hourendurance blocktest.From2010
onwards,inorderto leverage itsknow-how,the companydiversifieditsactivities:the first
WESTT SOLUTIONStestbenchwasinstalledin2011 andthe firstR&T projectwassignedin
2012. Since then,the DGEN programhas undergone more than2,000 cycles,1,500 hours of
operationsandtwosuccessful 150-hourendurance blocktests.DGEN enginesare regularly
producedforboth the developmentof the programand the WESTT SOLUTIONSproductfamily.
DGEN 380 Engine Cutaway 1
77. 4.2.2. Garmin
Garmin'smission isto be an enduringcompanybycreatingsuperiorproducts for
automotive,aviation,marine,outdoor,andsportsthatare an essential partof our customers’
lives. Garmin'svisionistobe the global leaderineverymarket,andthe productswill be sought
afterfor theircompellingdesign,superiorquality,andbestvalue. The foundationof Garmin's
culture ishonesty,integrity,andrespectforassociates,customers,andbusinesspartners. These
3 words"BuildtoLast" describe the products,company,culture andthe future.Asa leading
worldwide providerof navigation,Garminiscommittedtomakingsuperiorproductsfor
automotive,aviation,marine,outdoorandfitnessmarketsthatare an essential partof our
customers’lives.
Garmin'svertical integrationbusinessmodel keepsall design,manufacturing,marketing
and warehouse processesin-house,givingthemmore control overtimelines,qualityand
service.Theiruser-friendlyproductsare notonlysoughtafterfortheircompellingdesign,
superiorqualityandbestvalue,buttheyalsohave innovativefeaturesthatenhance the livesof
the customers.
DGEN 380 Flow Visualization 1
78. Garmin G1000®
The Standard inGlass FlightDeckCapability
Certifiedonabroadrange of aircraft models
Integratesvirtuallyall avionics
See clearlyeveninIFRconditionswithSVT™
GFC 700 digital autopilot integration
The G1000 isan all-glassavionicssuite designedforOEMor customretrofitinstallationona
range of businessaircraft.Itisa seamlesslyintegratedpackage thatmakesflightinformation
easiertoscan and process.Itsrevolutionarydesign bringsnew levelsof situationalawareness,
simplicityandsafetytothe cockpit.
The G1000 putsa wealthof flight-critical dataat a pilot'sfingertips.Itsglassflightdeck
presentsflightinstrumentation,navigation,weather,terrain,trafficand engine dataonlarge-
format,high-resolutiondisplays.Itfeaturesaflexibledesign, G1000 adapts to a broad range of
aircraft models.Itcanbe configuredasa 2-displayor3-displaysystem, withachoice of 10" or
12" flat-panel LCDsinterchangeable foruse aseitheraprimaryflightdisplay(PFD) ormulti-
79. functiondisplay(MFD).Anoptional 15"screenisalsoavailable forevenlargerformatMFD
configurations.
The G1000 replacestraditionalmechanicalgyroscopicflightinstrumentswithsuper-
reliable GRS77Attitude andHeadingReference System(AHRS).AHRSprovidesaccurate,digital
outputand referencingof youraircraftposition,rate,vectorandaccelerationdata.It’seven
able to restartand properlyreference itself while youraircraftismoving. The G1000 also
includesthe GFC700, the firstentirelynew autopilotdesignedandcertifiedforthe 21st century.
The GFC 700 is capable of usingall data available toG1000 to navigate,includingthe abilityto
maintainairspeedreferencesandoptimize performanceoverthe entire airspeedenvelope.
4.2.3. Rockwell Collins
Rockwell Collinsisapioneerinthe design,productionandsupportof innovative
solutionsfortheircustomersinaerospace anddefense.Rockwell'sexpertiseinflight-deck
avionics,cabinelectronics,missioncommunications,informationmanagementandsimulation
and trainingisstrengthenedbytheirglobal service andsupportnetworkspanning150
countries.Workingtogether,theirglobal teamof nearly20,000 employeessharesavisionto
create the most trustedsource of communicationandaviationelectronicssolutions.
Rockwell'saviationelectronicssystemsandproductsare installedinthe flightdecksof
nearlyeveryairtransportaircraftin the world.Theircommunicationsystemstransmitnearly70
percentof U.S. and alliedmilitaryairbornecommunications.Whetherdevelopingnew
technologytoenable network-centricoperationsforthe military,deliveringintegrated
electronicsolutionsfornewcommercialaircraftorprovidingalevel of service andsupportthat
increasesreliabilityandlowersoperational costsforourcustomersthroughoutthe world,
deliverontheircommitments.
80. Rockwell Collinsisaleadingproviderof flightcontrol andnavigationsolutionsfor
commercial,militaryandUnmannedAircraftSystems(UAS).Theirflightcontrol systems
expertiseincludesautopilot,actuation,fly-by-wire,pilotcontrols,andenginecontrollers.The
flightcontrol productsexemplifyourcapabilitiesinsystemsengineering,precisionmachining,
fabrication,andassemblyof close-toleranceflightcritical partstomeetdesignandcertification
requirements.Regardlessof asystem’scomplexity,theirflightcontrolsensurethe stabilityand
safetyof flightoperation.
Fly-by-wire systemsreduceweight,improve reliability,andincreaseaircraftfuel
efficiency.Rockwell'sfly-by-wire systemshelpcreate afamiliarenvironmentforpilotsby
combiningcomputersoftware andhardware toemulate the lookandfeelof mechanical pilot
control systems.Movementsof the column,wheel,andpedalsare convertedtoelectronic
signalsandtransmittedelectronicallybywirestothe control surfaces.
4.2.4. Heroux-Devtek
Héroux-DevtekInc.isaCanadiancompany specializinginthe design,development,
manufacture,integration,testingandrepairandoverhaul of landinggearandactuationsystems
and componentsforthe Aerospace market.The Corporationisthe thirdlargestlandinggear
companyworldwide,supplying boththe commercial andmilitarysectorsof the Aerospace
market.The Corporationalsomanufactureshydraulicsystems,fluidfiltrationsystems,electronic
enclosures,heatexchangersandcabinetsforsuppliersof airborne radar,electro-opticsystems
and aircraftcontrols. The Corporation’semphasisonResearch&Development,itssystems
integrationaccomplishments,anditsengineeringprowessare increasinglymakingHéroux-
Devteka preferredpartnerforthe design,qualificationandmanufacture of completelanding
gear systems
81. 5. LEGAL and REGULATORY/ SAFETY
5.1. FAA Certification Strategy
Thissectionwill give averybrief overview of the aircraftandcomponentcertification
process. By no meansisthisto be utilizedasthe sole directionforthe process,butagenerality
for the purpose of understandingthe process.
In general,there are several phasesaccordingtothe FAA forthe entire aircraftapproval
process. The firstphase is todevelopthe conceptual design. The conceptual designwill consist
of the overall generalitiesof the aircraft;nospecifics.
Next,the requirements needtobe identified. The productdefinition,identificationof
associatedrisksanda mutual commitmenttomove forwardwiththose identifiedbyboththe
FAA and the applicantare completedinthisphase. Manymeetingstake place duringthisphase
and a preliminarycertificationboardmeetingisheld. Thisiswhere the proposedschedule for
the entire certificationprocessismade.
Next,the aircraftwill needtobe designedinaccordance withpropercompliances.
Specificprojectplanningisdone and a ProjectSpecificCertificationPlan(PSCP) ismade. Thisis
the FAA’sspecificcompliancesforwhattype of aircraftitis. For our purposes,we designedin
accordance withCFR§23, AirworthinessStandardsforNormal,Utility,AcrobaticandCommuter
CategoryAirplanes. Thissubchapterof the Federal AviationRegulationsforAviation
Maintenance Techniciansdefinesrequirementsof eachsubsystemandwhatkindof testingthey
mustundergo. Otherparts may be specifictolargersubcomponents,however. Anexample
wouldbe §33 talksabout the airworthinessstandardsof aircraftengines.
82. The implementationprocessiswhere youbegintosee results. The applicantmust
demonstrate theircompliance withthe FARAMTsubsectionsforthe particularsystems,show
compliance andcomformance tothe previouslyidentifiedrequirements,andhave afinal
certificationboardmeeting. Thisiswhenthe aircraftwill be inspectedandsafetyanalysiswill
be performed.
The final phase ispost-certification. Thisphase primarilydealswithprocessestoensure
continuedairworthinessstandardsare met. Thisincludescertificate managementforthe
remainderof the product’slife cycle.
To beginthe process,the applicantwill turninFAA Form8110-12 to the nearest
CertificationOffice,whichislocatedinChicago,IL. Initially,thisformwill be filledout
requestingatype certification.
The followingisageneral outlineof the entire process:
• Within2 weeksafter application:
• Acknowledgementof applicationissued
• FAA CertrificationProjectNotification(CPN) issued
• Within1 month after application:
• Projectteamidentified(FAA andApplicant)
• PreliminaryType CertificationBoardMeeting(PTCBM) scheduled
• Within1-3 months after PTCBM:
• ProposedCertificationBasisG-1issue paperpreparedandprocessingbegins(stage
1)
• PSCPdrafted
83. • Within4-6 months after PTCBM:
• Final CertificationBasisG-1issue paperclosed
• PSCPagreedandsigned,includingthe mutuallyagreedprojectschedule
• Within6-9 months after PTCBM:
• All issue papersclosed
• One monthprior to scheduledTC/STC/ProductionApproval issuance:
Compliance documentationsubmittalsshouldbe scheduledoverthe course of a
projectto be completedbythispointintime. More than onmonthmay be needed
insome cases,especiallywhensubmittalsare notFAA Designee approvedor
recommendedforapproval
The followingisidentificationof the keyplayersthroughoutthe processandtheirprimaryroles:
FAA and Applicant’sManagement –Providesacommitmenttothe Partnershipfora
SafetyPlanas well asprovidesleadershipandresources
FAA and Applicant’sProjectManagers – Jointlyorchestratesthe projectandappliesthe
PartnershipforSafetyPlanagreements
FAA StandardsStaff ProjectOfficers –Providesatimely,standardizedpolicyand
guidance
FAA and Applicant’sEngineersandDesignees –Applyregulationsandpolicytofind
compliance includingthe determinationof the adequacyof type designand
substantiationdata
FAA and Applicant’sInspectorsandDesignees –Determinesconformityand
airworthiness
FAA and Applicant’sFlightTestPilotsandDesignees –ConductsFAA flighttests
84. FAA Chief ScientificandTechnical Advisors(CSTA) –Providesexpertadvice andtechnical
assistance
FAA AircraftEvaluationGroup – Evaluatesconformance tooperationsandmaintenance
requirements
Belowisan example of the PSCPprocess:
85. 5.2. Risk Mitigation Strategy
Riskisa functionof likelihoodmultipliedbythe severity. Aslongasone of the variables
inthe functionisratedto be high,the projectwill be consideredtobe risky.The risk
managementapproachincludesfourphases,riskidentification,riskassessment,riskresponse
planning,andriskevaluation.
5.3. Risk Identification
In the firststage of riskmanagement,itiscrucial topinpointthe risksandfocuson
the risksthat are highlylikelytocause the projecttofail.Riskscan be foundinternallyand
externally.The internalrisksinclude marketrisk,assumptionrisks,andtechnical risks. The
project,like designinganewaircraftinvolvesaseriesof highrisksinthe marketandtechnical
aspects.
Withthe DGEN 380 engine,the new designisclassifiedasaVLJ or PLJ.The market
for these typesof aircraftisnot completelyexploited.Withthe freshnessof the market,the
definitionof the marketremainsambiguous.Inaddition,whenthe projectisdelegatedbythe
Price Induction,the requirementsfromthe customerneedtobe fullydefinedandincludethe
detailstothe mostextent.Failure todefine the marketorthe customer’srequirementsclearly
couldresultinthe risksof misleadingthe directionof the project.Withthe marketresearch
made by the projectteam,there are three modelscurrentlyonthe marketsold bythree
differentcompanies,butusingthe enginesproducedbythe same companyPratt& Whitney
(Pratt& Whitney,n.d.).However,avarietyof aircraft inthe same class are eitherunderflying
testor in the phrase of undergoingdevelopment.Beingunable tokeeptrackof the newly
introducedproductsandmodifyingthe new designcouldbane the competitivenessof the new
design.
86. Anothermajorrisksexistinginternallyinthe projectisfromtechnical aspect
includingfouressential features:maturity,complexity,quality,andconcurrencyof the project.
As a newlyformedteamthathasn’tdabbledinthe aircraftdesigningforalongtime,lackof
experience andknowledgecouldleadtothe more time consumingandmore expensive.With
the newlydeveloped engine,the innovationandcreativityinthe projectcanalsoincrease the
risks.Besides,the complexityof aprojectlike aircraftdesigningcanalsoaffectthe likelihood
and severityof the risks.The proceduresinaircraftdesignedare highlyrelated andinvolve
numerousinterrelations.The calculationsandestimationsonthe TOGWdictatesthe
calculationsof the restparametersmostly.The estimationof TOGWcan be influencedby
variousindustrial andeconomicfactorsbesidesthe technical factors,suchasthe customer’s
requirement,budgetof the project,manufacturingprocess,andetc.Like all of the otherdesign
projects,the end-itemof aircraftdesignistoproduce the aircraftdesignedbythe team.Inthe
processof the design,the end-itemcannotbe completelyproducedorfullytested.
Consequently,the extentof testabilityandproducibilityalsohave effectsonthe risksof the
project.Last butnot least,fromthe Gantt chart of the project,due to the time constrainonthe
project,several sequential activitiesoverlapeachotherandmostof the activitiesare dependent
on the otheractivities.
As forexternal risks,the projectislimitedtothe followingfactors:government
regulations,customerneedsandmarketconditions,material orlaborresources,andphysical
environment.Asahighlyregulatedindustry,the projectof designinganew aircrafthas to be
compliedwiththe FAA certificationsandtestingstandards.The amountof demandsandthe
conditionsof the marketforthe newproductcan alsoaffectthe successof the project. Lack of
materials,resources,laborforces,andterraincanalsohave an influence onthe risksof the
project.
87. 5.3.1. Risk Assessment
There are plentyof methodsof assessingthe levelsof risks.Since riskisafunctionof
twovariables,likelihoodandseverity.The equation(Eq.5.2.2-1) below presentsthe risk
function:
Risk = Likelihood × Severity--------------------- Eq.5.2.2-1
The methodof risk matrix will be usedforthisprojecttoevaluate the risks identifiedinsection
5.2.1. The likelihoodof ariskcan be assessedfromfive levels,veryunlikely,unlikely,possible,
likely,andverylikely.Likewise,the severityof ariskcan alsobe dividedintofivelevels,low,
minor,moderate,significant,andhigh.The matrix below presentthe resultof ariskconsidering
fromboth likelihoodandseverity:
Severity
Likelihood
Low Minor Moderate Significant High
Very Unlikely Low Low Med Low Medium Medium
Unlikely Low Med Low Med Low Medium Med High
Possible Low Med Low Medium Med High Med High
Likely Low Med Low Medium Med High High
Very Likely Med Low Medium Med High High High
Afterevaluatingeachriskidentifiedinthe firstphrase,risksare ratedas mediumhighandhigh
will be the onestobe focusedtodeal with.Those risksare fromthe conflictsamongmarket
conditions,productdemands,andthe technical concerns.
88. 5.3.2. Risk Response Planning and Reevaluation
Five thingscanbe arrangedto manage risks,transferring,avoiding,reducing,
acceptingrisks,andcontingencyplanning.Riskscanbe transferredbypurchasinginsurance and
by specifyingthe responsibilitiesandrisksof eachgroup. The groupsinvolvedinhighlyrisky
activitiesshouldbe constantlymonitoredbythe riskmanagementteamorhigherauthority.
The third methodof managingriskisto avoidrisk.However,avoidingrisksinahighlyperplex
projectlike aircraftdesigningcouldpotentiallyincrease the complicityof the project,which
contributestomore risks.Therefore,avoidingthe riskisnotrecommendedforthisproject.
Insteadof avoidingrisks,the riskscanbe reducedor mitigatedbyreducingthe
likelihoodandseverityof technical risks.Beforeputtingthe designintoproduction,modelsand
simulationsshouldbe formedandtestedtoimprove the performance of the design.
Additionally,the projectteamshouldalwaysconductaparallel developmenton the highlyrisky
tasksand assessthe performance of those tasksbefore proceedingtothe nextrelatedactivities.
Afterthe calculationonthe TOGW, the projectteamshouldcarefullyconsideraseriesof
conditionstorefine the sizingresultbefore usingthe resultforfurtherdecisionmade onother
parameters.Tocriticallyevaluate the projectbefore proceedingtothe nextone,the project
teamshouldhire some outside consultantstoassessthe project.Multiple contingencyplans
shouldalsobe proposedbasedonthe scenariosbrainstormedbythe projectteam.Throughout
the whole processof the project,the risksshouldalsobe monitored.Incase of new risksrising
up,the teamshouldinstall the contingencyplanassoonas the earlysymptomsof a risk show
up.Lastly,while estimatingthe budgetof the project,the financial teamshouldreserveapart of
budgetforprojectdelaying,costoverrun,andriskmanagement.
89. 6. PROGRAM MANAGEMENT
Program managementisthe processof managingmultiple related projectsatonce.
Where projectmanagementisoftenusedtodescribe one project,programmanagement
involvesmultiple projectsthatare all relatedandworkingtowardthe same goal or result.For
the KentAerocompany,there are many advantages of usingprogrammanagementtomanage
the separate projects thatgo into completinganentire aircraft,althoughitcanbe challenging
to pull off well.Issueslikegovernance andriskcanbe managedmore successfullyif asingle
teamis coordinatingefforts.
Changescan be managedmuch more effectivelyaswell.Completingall the relatedprojects
withinaprogram while stayingonbudgetandonschedule isfarmore likelywithgoodprogram
managementthanwithoutit. The three factorsthatdrive projectssuchas thisare performance,
schedule,andcost.
6.1. Modification or New System
Currently,there are no active planstomodifythe aircraft, add anynew systems or
components.However,the optionisalwaysopenaswe proceedintothe future. There isa
possible optioninthe future forentities,suchasthe governmenttopurchase thisaircraft,and
have it convertedtosuittheirneeds.Inthiscase,the rearpassengerseatscanbe removed,and
special equipmentcouldbe loadedandinstalledonboard.
6.2. Unique Program Circumstances
The unique circumstance forthisaircraftis the fact that we are buildingandthe designthe
airframe aroundtwoPrice InductionDGEN 380 turbofanengines,whichare veryefficienthigh
bypassturbofanengines,buttheyare still inthe experimental stage,andare notfullycertified
yet.The Flashmust alsogo througha detailedFAA certificationprocess,whichwasdescribedin
the previoussections.
