Group 6 PDR Document

8/4/2019 Group 6 PDR Document

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AAE 451Senior Aircraft Design

Preliminary Design Review4-27-06Prof. William A. Crossley

Group 6John Collins

Chad Davis

Chris Fles

Danny Sze Ling Lim

Justin Rohde

Ryan Schulz

Ronald Wong

Yusaku Yamashita


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Table of Contents

Executive Summary...........................................................2

Market Review....................................................................3Design Requirements........................................................4Carpet Plots .......................................................................6Aerodynamics.....................................................................10Structural Analysis............................................................20Component Weight Breakdown........................................23Stability...............................................................................27Configuration and Dimensions ........................................34Propulsion...........................................................................39Fuel Selection.....................................................................41Direct Operating Costs......................................................45Acquisition Cost.................................................................47Production Cost.................................................................48Concept Comparison.........................................................51Conclusion..........................................................................52References..........................................................................53Merit Pool............................................................................54

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Executive Summary

Current forecasts on the world fossil fuel supply have shown the possibility of nearing

peak oil production within the next few decades. To date, no significant steps have beentaken to prevent the effects of declining oil availability on the aviation community.However, a large portion of the world economy relies on affordable air transportation.Thus the need has arisen to design an aircraft which will remain reliable in a time oftransition away from oil dependence.

The goal of an aircraft design company is ultimately, like any other company, to beprofitable. It has been identified that opportunity for profit exists in the emerging air taxiand air charter markets, as well as in sales to private companies. Through extensivestudy, it has been determined that the best way to effectively capture a portion of thesemarkets is through the development of a light, single turboprop aircraft powered byalternative fuel the Yamasan 2006.

The Yamasan 2006 is capable of carrying six passengers and two crew members on a 600nautical mile trip. The aircraft is designed to access runways as short as 2,100 feet. Thisenables the owner/operator to avoid congestion at major airports, providing moreconvenient point-to-point service. The Yamasan 2006 cruise speed of 250 knots iscomparable to aircraft of similar function, and is capable of a maximum 263 knots.

The design process led to a configuration that is not typical within the general aviationmarket. The Yamasan 2006 is a pusher-type turboprop aircraft with a canardconfiguration and vertical stabilizers located at the outer end of each wings. The designhas been shown theoretically feasible - both aerodynamically and financially - to be

introduced into the desired market. It is a unique solution for both an uncertain economicfuture and the needs of emerging air transport services.

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Market Review

The target market for Yamasan 2006 consists primarily of air taxi services, air charter

operations, and corporate flight departments. The aircraft is also suitable for use in cargoand medical emergency roles. The expected market exists not only in the United States,but also to a large extent in Europe and Asia. The combination of an alternatively-fuelledaircraft, a broad market, and effective utilization of the many advanced design toolsavailable today, will allow the aircraft to be successful and competitive against currentand future fleets.

As of 2002, the global business aircraft fleet included 9,785 turboprop aircraft. Of theseaircraft, air taxi service Pogo estimates that a 2% capture of flights of less than 500nautical miles will be required to remain profitable. This is reinforced by a LinearAirestimate of 16,000 passengers per day traveling in this range. Furthermore, EclipseAviation projects air taxi services to account for approximately 20% of the entirebusiness aircraft fleet by 2015, whereas the current figure is approximately 3.9% (382aircraft.

Through the study of the market size and competition, customer attributes have beenidentified, a Quality Function Deployment (QFD) matrix employed, and trade studies performed to define the necessary design requirements of the Yamasan 2006. Theaircraft concept aims to deliver an affordable and time-saving means of travel to the end-user at affordable and competitive operating cost levels for the operators. It also aims toprovide a positive solution to potential global oil supply problems. With flexible interiorconfigurations, it can also provide cross-platform capability for different market needs.

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Design Requirements

Figure 1 - Design Mission

Design Mission

The anticipated design mission for Yamasan 2006 is given in Figure 1. Based on severalof the performance parameters and requirements, it outlines the segments of an idealoperation; including taxi time, landing, and emergency procedures. A 10-minute taxitime was determined using the assumption that the aircraft will be utilized primarily forsmall airport operations, and thus will not face significant delays. Based on FARrequirements, the aircraft will have a 45 minute fuel reserve in the event of anemergency, such that its range will be extended beyond 600 nautical miles if necessary.

This design mission was selected because of its wide use in the current general aviationmarket. The average commercial flight distance is approximately 500 nautical miles.The selected design mission, at 600 nautical miles, easily accommodates the majority ofthe current public market. The length of this mission will maximize the capabilities ofowner/operators, such as air taxi services, to make point-to-point flights available tocustomers who simply seek to bypass large airports and reach the same destination morequickly. Additionally, a load of six passengers was found to be average for the typical business flight in the selected range, which requires two flight crew members on allflights.

Requirement Selection

The takeoff distance (ground roll) was chosen in response to the capabilities of similaraircraft currently in the market. The distance of 2,100 feet is shorter than that of themajority of the competition; this allows the Yamasan 2006 to access a greater number ofpublic airports, thereby increasing its utilization potential and creating an advantage overmany competing aircraft. The cruise speed is slightly higher than the primary competingaircraft, though it does not exceed the capabilities necessary for this regime of flight.

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The acquisition and direct operating cost requirements are based on a regression ofcompeting aircraft. The goal of the Yamasan 2006 design was keep the costs at equal orlower levels, without detracting from the performance and capabilities of the aircraftitself. The acquisition cost of the Yamasan 2006 is $1.725 million slightly lower thanthe design requirement. The direct operating cost of $450 per flight hour represents asignificant improvement over the $550 per hour requirement. The results of Yamasan2006 versus the initial requirements are given in Table 1 below.

Yamasan 2006 Design Requirement

Payload (lb.) 1500 1500

T.O./Landing Distance(Ground Roll, ft)

2100 2100

Capacity 6 passengers, 2 crew 6 passengers, 2 crew

Range (nm) 600 w/200 nm divert45 min. loiter

600

Speed (kts.)250 (Cruise)265 (Max.)

250

Acquisition Cost ($M) 1.725 1.8

D.O.C. ($/hr.) 450 550

Table 1 - Comparison of Yamasan 2006 with Design Requirements

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Carpet Plots

In order to determine aircraft characteristics such as wing loading, power loading, andaspect ratio, carpet plots were used. The carpet plots also involve better constraints than

in previous analyses because the models for performance are more accurate. The plotsinclude aircraft gross take-off weight as a function of wing loading, power loading, andaspect ratio; also, the carpet plots allow the aircraft characteristics that result in the lowestgross weight to be determined. Software programs such as FLOPS were utilized, withaircraft characteristics and mission requirements as input. Parameters such as take-offfield length, maximum cruise speed, and aircraft gross weight were calculated. However,it was decided to develop a program which could be tailored for this specific aircraft typeand design mission. The new program greater control and understanding of the carpetplots.

The program developed for the purposes of this design is comprised of three main parts.The first part is the relationship between wing loading, power loading, aspect ratio, grossweight, velocity, and the empty weight fraction. The historical regression presented inTable 6.2 (Raymer1) was used. This function represents the empty weight fraction thatwas used in the second part of the program, which is the mission weight fraction analysis.The weight fraction analysis is based on the fuel burn for mission segments, and Equation6.12 (Raymer) shows how the cruise segment weight fraction is analyzed. Similarequations are used to for the other mission segments, such as climb and loiter. Theprogram uses the empty weight fraction, crew and passenger weights, fuel fractionanalysis, an initial gross weight estimate, and an iterative process used in conjunctionwith Equation 6.1 (Raymer). The process continues until the initial gross weight estimateequals the calculated gross weight. The program determines the aircraft gross weights fora wide range of wing loadings, power loadings, and aspect ratios.

Within the first two parts of the program, other aircraft characteristics or missionparameters are not fixed by design criteria. Some aircraft characteristics, such as parasitedrag or thrust specific fuel consumption, are determined through the aerodynamicanalysis and engine models presented in this report. Other flight parameters such as thelift-to-drag ratio for cruise are determined using functions similar to Equation 6.13(Raymer). One mission parameter that is not constrained is the cruise altitude. There isno design requirement for cruise altitude, so it was necessary to calculate the altitude thatwould maximize the aircraft efficiency while maintaining the cruise speed designrequirement. Equations 5.13, 17.25, and 17.28 (Raymer) represent the flight conditionsfor best cruise range. The constraint equation for cruise speed represents the maximumcruise speed based on the values for wing loading and power loading. With fixedcharacteristics, such as wing loading and power loading, the main variable in theseequations becomes the dynamic pressure. The variation of altitude, and how it affectsbest cruise speed and maximum cruise speed was studied. The best cruise speed and amaximum cruise speed based on a cruise power setting have intersecting trends whenplotted against altitude, which can be seen in2. The intersection occurs at an altitude of 20,500 feet, which represents the best cruisealtitude and the altitude at which the design mission cruise segment is flown the weightfraction analysis.

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15000 16000 17000 18000 19000 20000 21000 22000 23000 24000 25000310

320

330

340

350

360

370

380

390

400

410Velocity vs. Altitude

Altitude (ft)

Velocity(ft/sec)

Best Cruise

Cruise Power Selected Cruise Altitude

Figure 2 - Cruise Velocity vs. Altitude

The third part of the carpet plot program is the constraint analysis. The two most strict

design constraints for the aircraft based on the initial constraint diagrams were the take-off distance and cruise speed. Take-off distance and cruise speed are based on the sameparameters that affect the aircraft gross weight, including the wing loading and powerloading. When the program calculates the aircraft gross weight corresponding to a givenwing loading, power loading, and aspect ratio, it also calculates the maximum cruisespeed and take-off distance of that aircraft. Using the aircraft parameters and weightsthat meet the requirements, the trends of the maximum cruise speed and take-offconstraints can be plotted along with the aircraft gross weight, wing loading, and powerloading.

The team uses the three main parts of the program to determine the minimum grossweight that meets the design requirements as well as the corresponding aircraft

characteristics. The carpet plots shown below in Figures 3, 4 and 5 are used to determinethe best aspect ratio, shown in Figure 6. Using the carpet plot analysis, the minimumaircraft gross weight was determined as 6500 pounds at an aspect ratio of 7.6, a wingloading of 32 lb/ft2, and a power loading of 0.1632 hp/lb.

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7500

8000

9000

h

t(lbs)

Figure 3 - Carpet Plot (AR = 6.08)

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7000

7500

8500

h

t(lbs)

Figure 4 - Carpet Plot (AR = 7.6)

9000

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Figure 5 -Carpet Plot (AR = 9.12)

Figure 6 - Carpet Plot (GTOW vs. AR)

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Aerodynamics

Wing Airfoil Selection

NACA 44 series and 5 digit series airfoils were considered as possible candidates for thewing airfoil. NACA airfoils were chosen because of experimental data were easilyavailable. Another benefit to using NACA airfoils was that many programs such asXFOIL, XFLR5 and Java foil were able to give good prediction with the empirical datafor many NACA airfoils and thus the aerodynamics performance could be accuratelypredicted. Another advantage of the NACA 44s and 5s series airfoil was that surfaceroughness has little impact on the lift properties2. Thus, the airfoil would not require anykind of expensive surface finishing which may contribute to additional production costs.

As the canard interaction with the main wing during takeoff is difficult to obtain withoutany kind of Computational Fluid Dynamics (CFD) analysis or Wind tunnel

experimentation verification, it was decided that at this current stage, the wing would berequired to carry all the lift during takeoff. Based on our constrain diagrams, the max liftcoefficient Clmax required would be 1.5 at takeoff.

The other aerodynamic properties that are important in the selection process are: lowsectional drag coefficient Cd, high L/D ratio at cruise and low pitching momentcoefficient Cm. The most important characteristics for Yamasan 2006 would be to meetthe Clmax requirement and at the same time induce the lowest drag at cruise as possible,thereby achieving a high L/D ratio.

In Abbott and Doenhoff3, empirical data for the lift and drag coefficients of variousNACA airfoils have published. Five NACA airfoils were identified as suitable airfoils

for Yamasan 2006. They were NACA 4412, 4414, 22012, 23012 and 23014. Thicknessto chord ratio of at least 12% were chosen so that the wing would have adequate volumeto store the fuel and other mechanical or electrical control systems (hydraulics, etc).

An airfoil analysis program JavaFoil4 was then used to analyze the lift and dragcoefficients of these five airfoils. JavaFoil uses the same multi-panel potential flowmethod used in XFOIL, and the results obtained were comparable to XFOIL andempirical data as long as flow separation did not occurr. The main advantage of JavaFoilis its ability to analyze the lift coefficient of a finite wing and flapped wing.

Figure 7 shows the lift coefficient at various angles of attack for the five airfoils at a

Reynolds number of four million. It was observed that the NACA 44 series airfoilgenerates a much higher lift coefficient than the NACA 5-digit series. However, with a20 degree flap deflection at the - chord location, all the airfoils could achieve Clmax of atleast 1.5 at approximately 5 degrees of rotation. Thus, the airfoil which induces the leastamount of drag would satisfy the aerodynamic design requirements of the Yamasan 2006.

Figure 7 Lift coefficient curve at Reynolds number of 4,000,000

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Figure 8 -Cd at various Reynolds number

Figure 8 shows the drag coefficient of the five NACA airfoils at various Reynoldsnumber. It can be seen that the NACA 22012 has the least drag coefficient among thefive airfoils. Thus NACA 22012 was selected as the wing airfoil.Figure 9 shows the lift coefficient curve for the NACA 22012 airfoil with and without thedeployment of flaps, at a Reynolds number of four million. It can be seen that thedeployment of flaps increased the lift coefficient significantly.

Figure 2 - Effect of Flaps deployment to Lift Coefficient at Takeoff

Figure 30 - Drag Polar

Figure 10 shows the drag polar of the NACA 22012 airfoil during takeoff and cruise. Itcan be seen that during takeoff the higher lift coefficient also increases the lift-induceddrag significantly, thus creating a higher Cd value.

Table 2 below summarizes the aerodynamic characteristics of the NACA 22012 airfoilfor the Yamasan 2006.

NACA 22012 Cl Cd Cm.25Takeoff (20 o Flaps and 5o Rotation) 1.622 0.12356 -0.207Cruise 0.22 0.00784 -0.007

Table 2 - Aerodynamic Characteristics of the NACA 22012 airfoil

Canard Airfoil Selection

The airfoil of the canard was selected based on the cruise L/D requirement for Yamasan2006. L/D can be calculated using equation 6.13 from Raymer, and is given by:

( )0

110.89

1

/

D

L

qC WD

W S S q Ae

= = +

Where CD0 = Parasite Drag at cruiseW/S = Wing loading at cruiseA = Aspect Ratioe = Oswald Efficiency Factor

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The actual L/D at cruise can be calculated by dividing the lift coefficient of the aircraftwith the drag coefficient at cruise. CL at cruise can be calculated by:

At Cruise,

11.77

wing canard

total wing canard

c L l l

w

L

D

L L L

CC C C

C

L C

D C

+

+

= =

Where Cc = Mean Aerodynamic Chord of the Canard

Cw = Mean Aerodynamic Chord of the WingCL at cruise = 0.320CD at cruise = 0.027

The drag coefficient CD at cruise can be obtained by using the component buildup methodas discussed in the drag section. NACA 2212 airfoil was selected, as it would provideadequate Cl at cruise and subsequently, an L/D ratio of 12.26.

Although the lift contribution of the canard during takeoff were ignored, a sufficientlylarge Cl is required during takeoff to rotate the aircraft. From the weights and balancesection of this report, lift coefficient of about 0.823 is required to rotate the aircraftduring the takeoff phase. In order to achieve this high Cl value, a flapped canard is

required.

Table 3 summarizes the aerodynamic properties of the NACA 2212 airfoil:

NACA 2212 Cl Cd Cm.25Takeoff (With 10 degree flap deflection) 0.823 0.10154 -0.137Cruise 0.437 0.03432 -0.039

Table 3 - Aerodynamic Characteristics of the NACA 2212 airfoil

Vertical Tail Selection

As the vertical tail on the Yamasan 2006 is located at the wing tips of the aircraft, thevertical tail also acts as a winglet. As the main reason for locating the vertical tail at thewing tip is to improve the lateral stability and rudder effectiveness of the aircraft, themain criteria for the airfoil selection is to ensure that the camber of the airfoil should begreater than that of the wing to produce enough side force and have a four degreeleading-edge-out incidence angle4. The NACA 4412 airfoil was chosen, as it fulfills therequirements stated above. A 12% thickness-to-chord ratio was selected to provideenough space to house the rudder control mechanisms.

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Drag Calculation

The drag on the aircraft comprises of the following 3 components:

The parasite drag, CD0 mainly due to the skin friction drag and a small contribution fromseparation pressure drag.Lift induced drag CDi due to the downwash created by the wing tip vortices when lift iscreated.Compressibility wave drag due to a shock wave when the wing is at some critical Machnumber where transonic flow is achieved.

Total Parasite Drag

The compressibility wave drag can be ignored in this case as Yamasan 2006 would not beflying at speeds close to the critical Mach number. The parasite drag can be obtainedusing the component buildup method as discussed in Raymer, where:

( )&o MISC L P

fc C C WETC

D D D

ref

C FF Q S C C C

S= + +

Where Cfc = Component Skin Friction CoefficientFFc = Component Form FactorQc = Component Interference Factor

Swetc = Component Wetted SurfaceSref= Wing Reference Area

The Skin Friction Coefficient can be calculated using the formula:

2.58 2 0.65

10

0.455

(log ) (1 0.144 )fC

R M=

+

Where R = Reynolds number M = Mach number

The Component Form Factor can be calculated using the formula:

For the Wing, Vertical Tail, Canard

4

0.18 0.280.61 100 1.34 (cos )( / )

m

m

t t FF M

x c c c

= + +

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Where (x/c)m = chord wise location of airfoil max thickness point(t/c) = thickness to chord ratioM = Mach numberAm = sweep of the max-thickness line

For the Fuselage

3

601

400

fFF

f

= + +

,

lf

d=

Where l = length of fuselage = 40 ftd = diameter of fuselage = 5 ft

Table 4 summarizes the Parasite drag of each component:

Wing Vertical Tail Canard FuselageCruise Takeof

fCruise Takeoff Cruise Takeof

fCruise Takeof

fCf 0.0030

4

0.00349 0.00356

1

0.00411

8

0.0032 0.00368 0.0022

6

0.00255

FF 1.978 1.3643 1.5689 1.1675Q 1.25 1.1 1.3 1

Swet 434.5576 ft2 41.1551 ft2 67.98 ft2 72.3942 ft2

Sref 212 ft2

Cdo 0.01542

0.01768 0.001038

0.0012 0.0021 0.00241 0.0009 0.00102

Table 4 Summary of the Parasite drag of each component

Miscellaneous Drag , Cdmisc

The upsweep angle of the fuselage creates some form of drag which can bepredicted by using the formula below:

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( )0

2.5

max3.83

/ 0.00258

upsweep

upsweep

d

ref

Du A

q

D qCS

=

= =

Where u = Upsweep Angle in Radians from the Fuselage CenterlineAmax = Maximum Cross-sectional Area of the FuselageSref= Wing Reference Area

During the takeoff phase, landing gears also contribute significantly to the parasite dragand could be estimated using table 12.5 in Raymer [1].

( )

( )0

/ 0.25 0.3

/3 0.00926

wt sgears

gears

d

ref

D q f f

D qC

S

= +

= =

Where f wt = Frontal Area of the wheel and tirefs = Frontal Area of the strut

The dimension of the landing gear strut is approximately 2.725 feet by 0.3 feet indiameter and of the wheel is 1.2 feet by 0.3 feet in diameter.

The total parasite drag can then be calculated by summing up the parasite drag of eachindividual component and the miscellaneous drag and the multiplying the values by 5%to account for the drag due to leakage and protuberance.

Total Parasite Drag (Cruise) Total Parasite Drag (Takeoff)0.023231 0.045393

The total drag on the aircraft can be obtained by summing the parasite drag contributions

on the aircraft together with the lift induced drag. The lift induced drag can be obtainedfrom the following equations:

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2

0.68

1

1.78(1 0.045 ) 0.64

id LC KC

KAe

e A

=

=

=

Where e = Oswald Span Efficiency MethodA = Aspect Ratio

The total drag, CD can then be calculated as follows:

0

2

D D LC C KC = +

The total drag coefficient during takeoff does not take into consideration of the lift-induced drag produced on the canard. However, since the Yamasan 2006 was able to takeoff with around 50% of the available horsepower, this additional drag will not affect thetakeoff capabilities of the aircraft.

Total Drag Coefficient (Cruise) Total Drag Coefficient (Takeoff)0.0276 0.1577

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Loading

V-n Diagram

Figure 11 - V-n diagram of Yamasan 2006

Figure 11shows the V-n diagram of the Yamasan 2006. The maximum load factor thatcould be attained by Yamasan 2006 is given by approximately 2.6 (see below) at acruising altitude of 20,000 feet and the minimum is at -1.

At straight and level flight,

max

max

/

L

L

Lift Weight L nW

qSCn

W

qCn

W S

==

=

=

Where q = dynamic pressureW/S = Wing loading

The stall speed can be obtained by using the above formula and substituting n as 1 tosolve for the velocity and is calculated to be 68 knots Equivalent Air Speed (EAS).

The maximum speed is typically 50% higher than the cruise speed4 and is given by 273Knots (EAS). To convert from EAS to True Air Speed (TAS) the following formula isused:

SLTAS EAS V V

=

Where = density at current altitudeSL = sea level density

The summary of the EAS speed at different load factors are given below:

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20

Vstall Vdive Vcruise VAOA68 Knots 273 Knots 182 Knots 115 Knots

Table5 -SummaryoftheEAS

speedatdifferentloadfactors


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Structural Analysis

An important aspect of the preliminary design is a structural analysis of the most

important structural member of an aircraft - the wing. The first step in performing thestructural analysis is determining the loading conditions present on the wing. Whileconventional wings typically have an elliptical lift distribution, the vertical surfaces at thewing tip and canards that alter the airflow near the wing root create a different loadingcondition. Conventional wing tips provide significantly reduced lift when compared tothe rest of the wing due to vortices that form at the tip. The vertical surfaces of thecurrent design would change the location of the vortices, and the wing near the tip wouldperform closer to an idealized infinite wing than the typical elliptical distribution. Thiswarrants using a trapezoidal load distribution on the wing because the lift near the wingtip would not reach zero. The presence of canards disrupts the flow of air in front ofwing root, and this could potentially add significant shear forces on the structuralsupports. Therefore, it was decided to approximate a larger distributed load on the wingroot than the average force that the rest of the wing experiences. A compressive force onthe wing due to a crosswind on the vertical surface is also included in the analysis, and adiagram of the loading conditions is shown in Figure 12 - Loading Conditions

Figure 12 - Loading Conditions

The load factor experienced at the wing root is 3.6 times the normal wing loading. Thehistorical loading factor for similar aircraft is around 2.5, so the higher factor applied tothis structure accounts for a safety factor of approximately 1.5. The load at the wing tipis 2.4 times the normal wing loading. A lateral force applied on the vertical stabilizerresults in a 1,000 lb compressive force on the wing. A torque of 1,000 lb.-in. was alsoexamined to test the twist resistance of the structural members in the wing.

The most critical aspect of the load on the wing is the resulting moment. The moment onthe wing creates the largest stress on the wing root and requires the most structuralmaterial in order to support the moment without failure. Figure 43 - Moment vs. WingLocation shows the moment values as a function of wing location. Additionally, theshear forces are calculated and are shown in Figure 5 - Shear vs. Wing Location.

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Figure 43 - Moment vs. Wing Location

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Figure 5 - Shear vs. Wing Location

The preliminary design of the structural members of the wing was approximated as a boxbeam in order to perform the analysis. In order to provide better structural support nearthe wing root where the shear force and moment have the greatest magnitude, the boxbeam was divided into two segments, one for the wing root and one for the outer segment

of the wing. This would allow for placing more structural support in important areaswhile reducing wasted material near the wing tip, where the forces are minimal.Breaking up the wing into more segments would further reduce the material used in lowstress areas near the wing tip. Figure 15 - Box Beam Design shows the box beam design.

Figure 15 - Box Beam Design

The chord length and airfoil thickness were used to determine the dimensions of the wingand wing root and to ensure that the base and height of the box beam fit within the wing.The box beam base is also 2 ft shorter than the chord lengths to allow for control surfacesand high lift devices near the trailing edge of the wing. The team determined thethickness of the box beam by using a program that would reveal the minimum thicknessnecessary to withstand the loading conditions present on the wing. The minimum

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thickness is based upon design requirements such as the maximum deflection, twist, andbuckling before the box beam fails. The maximum deflection of the wing before failureis 10 inches at the wing tip, and the maximum twist of the wing is less than 3 at the tip.The program also includes a study of different material types such as steel, aluminum,and composites, which allows for a weight calculation for the necessary structuralmaterial based on the density of each material.

The overall weight of the supporting structure in the wing is the most important result ofthe analysis. An aluminum metal matrix composite (MMC) supporting structure wouldhave an approximate weight of 220 lbs for each wing. This is less than an all-aluminumstructure, which would weigh 260 lbs, and significantly less than steel, which wouldweigh 450 lbs. While composites such as aluminum MMC cost more than conventionalmetal supporting structures to manufacture, they offer considerable weight savings for theaircraft.

Currently, the component weight of each wing should be 200 lbs. The wing weight based

on the preliminary structural analysis is higher than this value largely due to the boxbeam approximation used in the analysis. The box beam approximation results in excessmaterial being used in areas near the wing tip, and a more in-depth and optimal structuraldesign would be necessary to reduce the overall weight. However, based on thepreliminary analysis, achieving the desired ultimate loading factor and structural weightappears to be feasible.

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Component Weight Breakdown

The mission requirement set a gross takeoff weight (GTOW) of 6800 lb. This was then

modified by the sizing and mission segment analysis to a value of 6500 lb.

The following section will verify the value above with a component analysis. Thecomponent weights are calculated using individual characteristics and equations from amodel of general aviation by Raymer. The engine weight is modified with a regressiondatabase of 132 single turboprop engines according to the power requirement determinedby the carpet plots.

Summing up the airframe and different components placed the aircraft dry weight at 3322lb. The breakdown is shown by Table 6 below, categorized into different functionalgroups:

COMPONENTS

WEIGHT (lb)

COMPONENTSWEIGH

T (lb)

StructuresGroup

Equipment GroupWing 400Canard 96 Flight Controls 127Vertical Tail 34 Hydraulic 7

Fuselage 119 Avionics 300

Landing gear(Nose)

15 Electrical 516

Landing gear(Main)

115Air condition andanti-ice

337

Furnishings 331PropulsionGroup

Installed Engine 800Engine 475

Gearbox 175Propellers 150

Fuel Systems 125

Total Dry Weight 3322

Table 6 - Dry weight build-up

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This is then added onto the mission payload and fuel, in Table 7.

COMPONENTSWEIGHT

(lb)

Total Weight Empty 3322

Useful Load Group

Crew 300- pilot 150-co-pilot 150

Fuel 1665-usable 1580-trapped 85

Passengers (6 PAX) 1200

TAKEOFF GROSSWEIGHT

6487

Table 7 - Take-off Gross Weight build-up

The values above can be justified as follows: an average man weighs around 150 lb, sothe pilot and co-pilot each weighs 150lb without baggage. The passengers will beallowed around 50 lb of baggage each, totaling to 200lb per passenger of weight allotted.

The fuel was calculated using the range 600nm, plus the 200nm divert required by theFAA. The fuel burn for a typical turboprop engine is 50gal/hr. Traveling at 250 knots,

adding 18 minutes in total for climb, descend and taxiing, gives a total of 4 hours of fuelburned, or 215 gallons. An estimate of 5% trapped fuel is reasonable, giving us 226gallons in the tanks. With a specific gravity of 0.88 - an equivalent of 7.344 lb/gal, thetotal fuel weight becomes 1665 lbs.

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A better illustration is given Figures 16 and 17, depicting the percentage of eachcomponent as a percentage of the EDW and the GTOW, respectively.

27

Note: List in legend isarranged in an ascendingorder with regards to thepercentage of the EDW

Figure 16 - Component weight as a percentage of aircrafts dry weight

Figure 17 - Weight build-up as a percentage of GTOW

Component Weig

26%Component Tot(empty)

pilot + copilot

16%

2

Hydraulics

Landing gear (Nose)

Vertical Tail

Canard

Note: List in legend isarranged in an ascending orderwith regards to the percentageof the GTOW


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An analysis of market data demonstrates that the breakdown of component weightdistribution of our concept (Table 8) to be superior when compared to our competitors in

several areas. It is important to note, however, that the percentages are only initialestimates, and do not constitute the final design weights.

Yamasan 2006

EADSTBM

700 [5]

EADSTBM

850 [5]

PilatusPC-12

[6]

StarshipBeechcraft 2000A

[7]

Fuel 25% 25% 25% 28% 25%

Payload

23% 18% 18% 25% 33%

Empty 52% 57% 57% 47% 42%

Table 8 -Comparison of Functional Group Weights (as a percentage of GTOW) with competitors

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Stability

Cent e r of Gravity

The center of gravity of the aircraft moves with differing configurations. The nose of theaircraft has been chosen as the datum point. The travel of the center of gravity is given inFigure 18 for all possible weight configurations.

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Figure 18 - CG location with different configurations


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The different configurations are illustrated in Table 9 below:

Mission

# Configuration

Weight

(lb)

cg position from

datum point (ft)

1 dry 3322 22.07

2 dry+fuel 4987 22.64

3 dry+fuel+crew 5287 21.93

4 LANDING

a dry+crew 3622 21.08

b dry+crew+row1 4022 20.62

c dry+crew+row2 4022 20.99

d dry+crew+row3 4022 21.35

e dry+crew+row1+row2 4422 20.58

f dry+crew+row1+row3 4422 20.91

g dry+crew+row2+row3 4422 21.25

h dry+crew+row1+row2+row3 4822 20.85

5 TAKEOFF

a dry+crew+row1+fuel 5687 21.55

b dry+crew+row2+fuel 5687 21.81

c dry+crew+row3+fuel 5687 22.06

d dry+crew+row1+row2+fuel 6087 21.46

e dry+crew+row1+row3+fuel 6087 21.70

f dry+crew+row2+row3+fuel 6087 21.94

g dry+crew+row1+row2+row3+fuel(GTOW) 6487 21.60

Table 9 - Differing Configurations and corresponding CG locations

The furthest forward cg location is 20.58 ft from the nose (case 4e), traveling to 22.06 ftwith the furthest rearward operating configuration (case 5c), moving to the most rearposition of 22.64 ft possible (case 2). Case 2 is not an operating condition unless theaircraft is configured for UAV operations, and is only considered for the purposes offorward limit placement of the main landing gear.

The wing strake design is chosen as it allows a forward movement of the fuel, whichincreased the static margin to an acceptable level. Previous studies showed that the c.g.of the wing cannot be placed 28 ft behind the nose, as it reduced the static margin belowthe desired 10%.

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A breakdown of the components contribution to the CG is shown in Figure 19 below, andtabulated in Table 10:

The distances shown here are given in the table below:

Component Weight (lb) Moment arm Distance (ft)

Canard 96 d_canard 15.4

Avionics 300 d_avionic 15.2

Landing Gear Nose 15 d_lg_n 14.1

Crew 300 d_crew 12.6

Row 1 400 d_row1 6.2

Fuselage + Furnishings + electrical 966 d_fus+fur+elec 5.3

Row 2 400 d_row2 2.5

Fuel 1665 d_fuel 1.1

Row 3 400 d_row3 1.2

Aircraft Condition and Anti-ice and

Pressurization 337 d_a_c+a_i+p 1.9Wing and Landing Gear Main 515 d_wing+lg_m 3.0

Hydraulics + Flight Control + FuelSystems

259 d_hyd+flc+fsys 5.4

Installed Engine 800 d_engine 6.9

Vertical Tail 34 d_vtail 9.9

Table 20 - Moment arm of aircraft components

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Figure 19 - Component contribution to CG location

Distance from datum point in feet


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Longitudinal Stability

Static Margin

The FAA stability certification requirement of an aircraft is a positive static margin.General and business aviation aircraft have a range between 5-40% margin. By definingthe wing and canard aerofoil and placement, the neutral point can be calculated as:

Where xac = aerodynamic center of the wing;

Vc = canard volume coefficient is, given by:

a and ac are the 3-D lift curve slope of the aerofoil sections. These are converted from the2-D slope values obtained from XFLR5 for the NACA 2212 and 22012 airfoils:

c

is the mean aerodynamic chord of the wings, and is a function of taper ratio and root

chord length. Table 11 shows the characteristics and placement of the wing and canard:Table 31 - Characteristics and placement of the wing and canard

ca

aVxx

c

cacn+=

WW

ccc

Sc

SlV =

AR

a

aa

D

DD

2

23

1+=

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The center of gravity is shown to travel as configuration differs, or as fuel is burned.Therefore the two most severe operable cases were analyzed (furthest forward case 4e;and furthest back case 5c). Case 5c will be the limiting case for the static margin, andthe value is calculated to be 12.2%. Case 4e will be the limiting case for the liftcontribution of the canard with regards to its moment arm, and is calculated to be 45.7%.

The operating envelope is shown to have a reasonable static margin, therefore we proceedto check for the balance of forces and moments:

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Canard Main wing

leading wing x pos (ft) 5.00 22.31

AR 5.00 7.60

Span (ft) 19.36 40.14

S (ft2) 75.00 212.00

Swetted (ft2) 152.96 432.35

MAC (ft) 3.91 5.60

1/4 MAC (ft) 0.98 1.40

Ybar (ft) y pos of MAC from fuselage 4.56 8.60

ac pos (ft) 6.05 24.04

Volume Coefficient 1.14 -

thickness to chord 0.12 0.12

taper ratio 0.70 0.40

root chord length (ft) 4.56 7.55

tip chord length (ft) 3.19 3.02

sweep angle at 25% MAC 0.00 0.00

sweep angle at leading edge (radians) 0.04 0.06 (degrees) 2.02 3.23

a-2D 6.64 7.81

a-3D 4.67 5.89


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Lc

LWc

Mc

Mxc

xw

lc

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NoseTail


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Distance (ft)moment arm between canard ac and wing ac, lc 17.99wing ac from cg, xW 3.46canard ac from cg, xC 14.53

Table 12 - Moment arm distances

Force Balance for takeoff: Lc + Lw W 0

Moment Balance: Lxc M Mc Lcxc = 0

From these two equations, equilibrium for the aircraft can be found:

Lift: L = CLqS

Where q = 1/2V2

S = reference area of the componentCL = 3-D lift coefficient of the component

Moment: M = CMqSc

Where c = mean aerodynamic chord

CM = 3-D moment coefficient of the component

T akeoff

Two scenarios were considered during takeoff for the two cases (4e and 5c):

zero flap deflection

20 deg flap deflection

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Figure 6 - Moment Balance modelling


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The following conclusion is reached: case 5c can only allow takeoff with a 20 degrees offlap deflection, while case 4e can allow takeoff under both scenarios. The comparison isgiven in Table 13.

Takeoff Case 5c Case 4eScenario 1

Cl canard required 0.339 0.807Cl wing 1.171 1.171Total Lift Generated 5211 5879Takeoff weight 5687 4422

Scenario 2

Cl canard required 0.657 1.320Cl wing 1.611 1.611Total Lift Generated 7442 8389Takeoff weight 5687 4422

Table 4 - Takeoff cases for the two scenarios

For a 6 degree rotation of the aircraft, 10 degrees of flaps are deployed for the canard,raising the Cm to -0.137. Balancing the moments led to the resultant lift required to beplaced at 972 lb. The CL value of this aerofoil is also increased with the flap deploymentto 0.823, which gave a lift of 1175lb 200lb greater proving that the aircraft is capableof rotation.

C ruise

The cruise altitude is defined to be 20,500 ft at a cruising speed of 250 kts. The forcecomparisons at this condition are given in Table 14.

Cruise Case 5c Case 4e

Cl canard required 0.0757 0.165Cl wing 0.22 0.22Total Lift Generated 5901 6654Maximum cruising weight 5687 4422

Table 5 - Cruise casesThe moment balance indicates the aircraft has a capability in climbing at 20,500 ft. Theforce equilibrium is achieved through cruise climbs.

Lateral Stability

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The spin-recovery scenario is considered for the lateral stability. This is a comparison ofthe tail-damping power factor of that required to recover and that provided by the verticaltail.

The tail damping power factor (TDPF) available is calculated using the tail damping ratioand the unshielded rudder volume coefficient. These two parameters are influenced bythe reference and unshielded areas of the tail, the moment arm of the tail to the cg in thelateral direction, and the reference area of the wing. These equations can be found inRaymer. The TDPF required is obtained using the rudder alone recovery option fromFigure 16.32 in Raymer gave a value of 0.02.

The effective uncovered area has been estimated to be 5/8 th of the total reference area.This is a reasonable assumption, as the vertical tail is placed outboard from the fuselageon the wing tips, and therefore will be the last to stall in the event of a spin. This has beenconfirmed with the 60 deg requirement blanket that extends from the wings leading edgeas shown in figure 16.31 in Raymer [1].

Table 15 lists the properties of the vertical tail. As can be seen, the TDPF value of thevertical tail is 1.5 times more than that required for spin recovery.

.

Table 15 Vertical Tail Characteristics

37

Characteristics

Aerofoil Section NACA 4412

Area of Vertical Tail (bothsides)

26.48 ft2

Area of v tail (effective in spin) 16.55 ft2

Span (both sides) 126.0 ftroot chord length 36.00 fttip chord length 23.69 fttaper ratio 0.66MAC 30.27 ftAR 4.16

TDPF 0.0031TDPF required 0.0020


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Configuration and Dimensions

Figure 21 - Three Views of Aircraft

Figure 22 - Diagram of Fuel Tank, Flaps and Engine location

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The final dimensions of the Yamasan 2006 are shown in Figure 22. The placement ofvarious internal components can be seen in Figure 23.

Canard Justification:

The canard configuration offers potential advantage over traditional configurations:

Trimmed maximum lift coefficient for a canard is higher than the conventional airplane.By proper canard/wing layout design it is possible to achieve better trimmed lift to dragratios with a canard design.The canard must be designed such that it stalls before the wing. This way a stable pitch- break is obtained. Otherwise the wing is allowed to stall before the canard, anuncontrollable and sometimes violent pitch up motion can occur. To trim out the negativepitching moment associated with deployment of wing flap as shown in Figure 2, thecanard must be able to develop rather large lift coefficients itself. This can be done withthe introduction of flaps on the canard, by varying the sweep angle of the canard or byvarying the angle of incidence of the canard.

Note that the wing leading edge is given a very large sweep angle or strake, whichserves two purposes:

It provides volume for fuel to be carried close to the empty weight center of gravity.It serves to delay wing stall.

Most of the canard aircraft features a variable sweep canard. This is used to trim out thenegative pitching moment of the wing flaps.

The pusher-prop arrangement is a feature which seems to be increasing in popularity.Pusher configurations allow better laminar flow over the fuselage, thus providing anincrease in the upsweep angle at the tail, as well as a slight decrease in the drag over thefuselage itself. The design also leads to a quieter cabin despite its placement on thefuselage, as the majority of noise from the engine and propeller can be shielded from thepassengers and will propagate downstream. Another advantage is the reduction in adverseyaw since there is no prop-wash factor on the vertical tail.

Vertical Tail

The vertical tails on Yamasan 2006 is located at the wing tips also act as winglets. Thisdesign can improve the efficiency ofaircraftby lowering thelift-induced drag caused bywingtip vortices. It also increases the effective aspect ratio of a wing and increases theamount of lift the wings can generate. Control of yaw is assigned to these verticalstabilizers.

39
http://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Lift-induced_draghttp://en.wikipedia.org/wiki/Lift-induced_draghttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Aspect_ratio_(wing)http://en.wikipedia.org/wiki/Lift-induced_draghttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Aspect_ratio_(wing)http://en.wikipedia.org/wiki/Aircraft


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Landing Gear Position

The choice of landing gear layout depends on the size, type and configuration of aircraft.In this case, the tricycle configuration is used for easy ground handling, better takeoff

ground clearance of pusher engine propeller, and is the most commonly used landing geartype.

Advantages are that the leveled attitude of nose wheel aircraft makes easier to load andunload and provides better visibility out of the cockpit while taxiing. Also, shallowerangle of attack makes for a faster acceleration on take-off.

Disadvantages concerned are that the nose wheel has to take a greater load than that of atail wheel. Retractable mechanism is also preferred as the drag is greater for a tricycleconfiguration.

There are two geometric criteria that need to be considered in positioning the landing

gear struts (Figures 23 and 24).

Tip-over criteria.Ground Clearance Criteria.

Figure 23 - Tip-over Criterion

The lateral tip over angle is shown in 23. Angle of 35 degrees is the minimum allowed towith tricycle gear design. The main landing gear must be positioned behind the aft center

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of gravity (cg) location. The 25 degree angle shown in Figure 4 represents the usualrelation between main gear and the aft cg.

Figure 24 - Longitudinal Ground Clearance Criterion

Figure 24 also shows the required ground clearance angle. The lateral ground clearanceangle applies to both tricycle and tail-dragger configurations. The longitudinal groundclearance angle applies to tricycles only. Usually tail angle (longitudinal groundclearance angle) is around 10 to 15 degrees. With a pusher configuration, clearance of 3feet and 17 degrees is given to safely take-off and land on the runway.

Figure 25 - Landing Gear Overview

When the aircraft lands on the runway, the landing gear shock absorbers and tires arecompressed and deflected such that kinetic energy of descent is changed into a mixture ofthermal and potential energy.

Static Load (lb) Tire Size Tire Dim (in) Flat radius (in) Max. deflection (in)

700 5.00-4 13.25 3.6 3.02

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800 5.00-5 14.25 4.1 3.00

1100 6.00-6 17.51 4.5 4.25

Table 16 Typical tire characteristics of light aircraft

The data in Table 16 relate the static load per tire to size. The similarly-configuredBeechcraft Starship5 uses a 19.5 inch x 6.75 inch tire. Since the Yamasan 2006 is two-thirds the size of the Beech Starship, a tire size of 16 inch x 6 inch was chosen forYamasan 2006.

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Propulsion

To determine the size of the engine a rubber engine sizing approach was used. A

database of 132 engines was compiled, from which relations of horsepower-to-engine-weight, length, diameter, and specific fuel consumption were gathered. A powerregression equation was determined for each of the sizing relations as a function ofhorsepower. Using information from the carpet plots, the initial horsepower requirementwas entered into the formulas to get preliminary engine specifications. Table 17 showsthe equations and values.

Weight = 448.34 [lbs]

y = 1.8211x^0.7915

SFC = 0.552 [lb/shp-hr]

y = 0.8214x^-0.0573

Length = 56.64 [in]

y = 6.1832x^0.3184

Diameter = 25.36 [in]

y = 4.8598x^0.2375Table 17 -Engine Size Regression Equations

As the aerodynamic characteristics and specifications were determined and updated, thethrust and drag relationships were continually evaluated and kept in agreement such thataircraft would be capable of flight during best range cruise, maximum speed, and takeoff.

Engine power was adjusted as necessary to maintain these values. Simultaneously a propeller was selected for cruise based on the initial propulsion information andaerodynamics requirements using Hamilton Standard propeller efficiency charts. As thedesign evolved the propeller was re-evaluated and reselected as needed to maintainadequate thrust in all flight regimes. Propeller helical tip speed was monitored to ensuretip speed was within acceptable limits throughout the entire propulsion system design andselection process.

Once all values were finalized, the engine specifications were taken and then compared toall 132 engines in the database. An existing engine was selected that matched therequired rubber engine generated specifications. The decision to use an off-the-shelfengine was agreed upon to minimize development cost of the aircraft. The engineselected is a Pratt & Whitney Canada PT6A-60A turboprop engine producing 1,050 eshpand weighs approximately 475 lbs (dry).

The pusher design of the aircraft placed strict limits on the diameter of the propeller toallow for adequate rotation on take off. The diameter was determined to be 8 ft from thedesign drawings. Due to this strict and specific limit, other engine and propellercharacteristics were determined based on the maximum diameter. Circular blade tipspeed was kept to Mach 0.75 for noise and efficiency reasons. From this the maximum

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rotational speed of the propeller was calculated to be 2,000 rpm. The forward motion ofthe aircraft was factored in to the tip speed to ensure the propeller tips remained subsonic;this helical blade tip speed is Mach 0.849.

Using propeller efficiency charts, or propeller thumbprints, from Hamilton Standard apropeller was selected for the aircraft. Propeller thumbprints are efficiency charts forvariable pitch propellers. A variable pitch propeller was chosen due to the large performance advantage over a fixed pith prop. The flight envelope of this aircraftdemanded a variable pitch prop to effectively perform the mission. Another aspect of thedesign mission is to allow landing at airports with runways of 2,100 feet for less. Toensure that a full stop can be safely achieved, reversible pitch is another feature requiredfor our propeller. Composite propeller blades were selected to save weight, which iscritical for weight and balance of a pusher setup, and also because composite blades havea longer FAA certified life due to the fact they are more durable than metal blades.

As stated previously, the propeller was selected to provide adequate amounts of thrust in

all flight regimes. This was done by analyzing the take off condition and maximumcruising condition. The advance ratio, J, was calculate for take off and maximum cruise,and the power coefficient, cP, was calculated for maximum power settings. Both 3-bladed and 4-bladed set-ups were investigated, and a 4-bladed propeller was selected. A4-bladed design was selected because it provided better overall efficiency at max speed,and its takeoff efficiency was also very favorable. The propeller that was ultimatelyselected is a Hamilton Standard 4-bladed, variable pitch propeller with a design c L of0.500 and an activity factor of 80. Table 18 below shows the propeller efficiencies:

cp = 0.2005 P (%) (deg.)J = 0.5155 47 27

J = 1.6655 90.5 38Table 18 Propeller Efficiency

The design mission requirement of a take off ground roll of less than 2,100 feet was usedto double check the propulsion sizing. The ground and balanced field length of theaircraft was calculated using equations 17.101-103 and 17.112-114 found in Raymer.These calculated ground roll value and balanced field length are 1,150 ft and 1,750 ft,respectively. The ground roll value is well within the design requirement - well enoughthat a 50 ft obstacle can be cleared is a distance less than the designed ground roll. Thiswas achievable because of the relatively high power-to-weight ratio of the aircraft.

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Fuel Selection

Alternative Fuels

Originally, the properties of several different alternative fuels for use in twin piston,single turboprop, and twin turbofan driven aircraft were examined. The advantages anddisadvantages of the best choices are discussed below.

Ethanol Blend - AGE85

AGE85 (Aviation Grade Ethanol) is a well-established and tested ethanol-basedalternative fuel. It is specifically blended for cold starting, which makes it an ideal fuelfor use in aviation where fuel line and carburetor icing must be avoided. It also burns

much cleaner than traditional aviation fuels.

The major disadvantage of AGE85 and other ethanol-based fuels is the dangerous effectson fuel system components. Ethanol may react with seals or lines, causing corrosion.Significant modification of these parts would be necessary if an ethanol-based fuel wereto be used. Furthermore, ethanol has a lower energy density than traditional aviationfuels. Lastly, the research with AGE85 has been primarily focused on driving pistonaircraft performance.

Pure Biodiesel B100

Biodiesel is made mostly from soybean oils, and contains no petroleum products.Because of this, availability and affordability of biodiesel would not be directly affectedby the world-wide petroleum markets. This is an important economic advantage in asituation where petroleum becomes prohibitively expensive. B100 already has asignificant production infrastructure, so the availability of the fuel would not be inquestion. Other advantages of biodiesel include the fact that diesel reduces engine wear,is comparatively safe to store and transport, and has some of the lowest harmfulemissions of any of the alternative fuels studied.

There are two major disadvantages of B100. The first is its high freezing and cloudpoints, as seen in Table 19. This causes problems for both operation and shipping of thefuel. Electric heaters would be necessary on the fuel tanks and engines during storage ofthe aircraft, and the shipping costs of the fuel would be significantly higher in cold-weather climates. The other major disadvantage is the reduced energy content. As Table20 shows, the energy density of B100 is almost 18% lower than traditional aviation fuels.Beyond the high freezing point and cloud point, researchers have discovered a problemwith using soybean-based biodeisel due to the limitations in growing soybeans fastenough to produce the required fuel needed for so many flights.

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No. 2 Diesel (petrol) B100 (pure biodeisel)Cloud Point (F) -9 35Pour Point (F) -17 32

Table 19 Pure Biodiesel Cloud & Pour Points

Jet-A Avgas No. 2 Diesel(petrol)

B100 (purebiodeisel)

Heat of Combustion (Btu/gal) 123099 115480 131295 117093Density (lb/gal) 6.676 6.092 7.079 7.328energy density by mass(Btu/lb)

18439 18956 18547 15979

Table 20 Energy densities of various fuel options

Biodeisel Blend B20 & others

B20 and other like fuels are blends of petroleum-based diesel fuels with pure biodeisel.Blending biodesiel with petroleum-based diesel relieves some of the problems with purebiodeisel. The freezing point is lower, and the energy density is higher than biodeiselmaking it a more feasible choice. However, the production of B20 still involves mostlypetroleum. Therefore, B20 can not be considered a renewable fuel, and the price of thefuel would change with the petroleum market, which would be a major disadvantagewhen petroleum is expensive.

BioJet Fuel

The University of South Dakota is currently developing a bio-matter based fuel that hasvery desirable properties. Freezing is not a problem for this fuel, as it operates normallydown to -75 F. It has very low emissions, and little modification would be needed to runin current turboprop engines. This BioJet fuel would be the ideal choice for analternatively-fueled aircraft. However, this fuel is only in the experimental phase, and noinfrastructure for the production and distribution exists. Designing an aircraft aroundthese obstacles would be an extremely large risk.

Choice of Fuel

Additional advantages and disadvantages of alternative fuels are given in Table 21. Thefuel chosen for this project was B100: 100% biodeisel. The advantages of the fuel interms of its environmental properties and its 100% renewability make it far superior tothe other fuels, despite its questionable cold start properties. B100 is relatively safe fortransport to airports throughout the country. Limiting factors include the time required toreproduce soybeans, corn, and other biomatter used in the refining process for productionof B100 fuel.

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Table 21 - Pros & Cons of Alternative Fuels

Due to BioJet fuel still undergoing developmental and experimental testing, deciding toallow the use of 100% Biodiesel to fuel our aircraft has caused for slight modifications tothe fuel tanks and powerplant configurations. To support biodiesel fuel, the addition offilters, heat exchangers, valve tanks, microprocessors, pumps, in-line fuel temperaturegauges, exchanger gauges, low fuel warning lights, visi-fuel & Vo alerts, buzzers, &switches must be calculated into parameters such as our cost and GTOW. Theconversion for each fuel tank will cost $10,000-$15,000 and weigh approximately 25-30lbs. This needs to be considered when calculating the weight of the fuel systems in order

to properly calculate the center of gravity for stability of the aircraft. Figure 26 belowshows the items needed for full conversion to Biodiesel from a standard petroleum fuelconfiguration:

Figure 26: Fuel system accesories

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TSFC and BSFC have been calculated for the selected fuel using ONX Parametric CyclicAnalysis software. Significant combination fuel and engine parameters are listed inTable 22:

Fuel Heating Value 16,000 (Btu/lb)Specific gravity: 0.87-0.89Efficiency: 80.55%Petroleum Efficiency: 83.28%TSFC: .567 lb/hr BSFC: .26-.30 lb/hr

Table 22 Overview of selected fuel parameters:

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Direct Operating Costs

The direct operating costs were determined using the equations in Raymer. These

equations are based on statistical data for historical aircraft. These equations predicted adirect operating cost of $450 per flight hour for the Yamasan 2006, which is considerablylower than our design target of $550 per flight hour.

The different components of the direct operating costs were calculated using theequations below (Raymer). The fuel costs were calculated by use of simple fuelconsumption calculations using the heating values associated with the fuel chosen for thisaircraft.

Crew:

12210

51

3.0

5+

oc

WV

Parts:

e

ea NCC

++

1910

582.1010

3.366

Labor:

mCFH

MMH

Where Vc = cruse speed (knots)W0 = empty weight (lbs)Ca = aircraft costCe = engine costNe = number of enginesMMH/FH = maintenance man hours per flight hourCm = maintenance cost per hour

Figure 27 - Breakdown of DOC ($/FH)

Figure 27 shows the breakdown of each category of direct operating cost. Indirectoperating costs such as vehicle depreciation, landing fees, fuel surcharges, and othertaxes were not included in the direct operating cost calculation. The amount of these costsvaries quite significantly with the location and use of the aircraft and therefore is left forthe customer to calculate. Table 23 compares the direct operating costs of the Yamasan2006 to that of other competing aircraft.

Aircraft DOC ($/FH)

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Baron G58 288Adam A500 450

Pilatus PC-12 400Yamasan 2006 450

Table 23 - Direct Operating Costs

The chosen concept is competitive in direct operating costs. It is important to note that asconventional fuel prices continue to increase, the Yamasan 2006 will see only a modestrise in DOC based on increased demand for alternative fuels while other conventionallyfueled aircraft will be forced to endure very significant increases in DOC. In thisscenario, the Yamasan 2006 will have a strong cost advantage for potential customers.

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Acquisition Cost

Once the aircraft gross weight is calculated, the aircraft acquisition cost can be estimated

in conjunction with the aircraft performance characteristics. Aircraft acquisition cost can be determined with a linear regression similar to the empty weight fraction. TheYamasan 2006 is not significantly different from current aircraft in the market, evenconsidering the alternative fuel usage. The fuel that was selected has similar properties topetroleum based fuels, and the turboprop engine selected does not need significantalterations to be compatible. The general similarities of the concept and the competitorsshould indicate that the acquisition cost is accurate. The overall acquisition cost for theconcept aircraft calculated using the cost model and regression is $1.725 million. Inaddition to the calculation of a single acquisition cost, in previous portions of the projecta trade study was conducted, and the variation of mission parameters and aircraftcharacteristics and their effects on aircraft weight and acquisition cost were studied. Anexample of that trade study can be seen in Figure 78 - Trade Study (Range and Speed).

Regression Cost Model

( )DexpVRwCost CBA0 =

2.5

Figure 78 - Trade Study (Range and Speed)

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Production Cost

The projected cost to produce a new model of aircraft can be difficult to estimate,

particularly for a startup company. Manufacturers with significant, long-term experiencein aircraft production (i.e. Cessna, Boeing, and Airbus) employ proprietary models topredict the costs of producing new aircraft. These models have proven extremely reliabledue to relatively unchanged procedures in engineering, tooling and manufacturing processes between the production of new and previous aircraft lines. A new aircraftcompany, however, has no previous production experience on which to base costestimates. Therefore, industry-averaged regression models must be used as guidelines forthe development of a new aircraft.

For the purposes of this project, the DAPCA (Development and Procurement Costs ofAircraft) IV model suggested by Raymer was used to estimate production costs. TheDAPCA IV model, developed by the Rand Corporation, is based on CERs (CostEstimating Relationships) which are averaged from the entire aircraft industry. Whilethis model provides a good overall trend in the dynamics of cost versus number ofaircraft produced, inaccuracies arise based on the size and capabilities of a particularaircraft manufacturer.

In order to account for these factors, the DAPCA IV model was studied in relation toexisting production costs for light aircraft. This allowed for multipliers (fudge factors)to be determined to compensate for the size and material usage of a new line of smallaircraft. Such factors were also used to compensate for the use of composite materials inthe aircraft structure.

Two primary sources were used against which to benchmark the DAPCA IV costanalysis against that of the Yamasan 2006. First, the Meyers Aircraft Company providesan analysis6 of two light aircraft of different empty weights using a modified version ofthe DAPCA IV model. Both analyses were performed at production numbers of one andfive-hundred aircraft. Secondly, a NASA airframe cost model7 was used with varying production numbers. In the NASA model, the term airframe encompasses allmanufacturing, production, and installation costs, but does not include developmentsupport (variable) and inventory (fixed per unit) costs.

The resulting production cost model is given in the MATLAB script mfgcost.m (seeGroup 6 files on course website) This DAPCA IV-based model has been adjusted (usingthe previously mentioned factors) to accurately reflect the actual production costs ofaircraft with empty weight similar to that of the Yamasan 2006.

The projected production cost of Yamasan 2006 per number produced is given in Figure29. Initial production is three aircraft, including two flight test units. At this point in theproduction, the total program cost will reach approximately $15 million. As productionis increased, the steady-state cost approaches a value of $1.38 million per unit. At aprofit of 25% for the initial production period, the acquisition cost becomes $1.725million, as previously determined. Assuming a constant profit margin during the initial

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years of production (estimated previously at 50-100 aircraft per year), the cost would berecovered after the sale of 316 aircraft (Figure 30).

Many of the costs to develop the new aircraft will remain constant during production,regardless of the number of aircraft produced. These fixed costs include the engine,avionics, materials, number of flight test aircraft, and inventory. Several costcomponents, however, will decrease in relation to the applied learning curve. Thesevariable costs include engineering, development support, tooling, manufacturing, andquality control. As production increases, the cost per unit of variable-cost componentswill decrease and eventually reach a steady-state value. Figure 31 shows the percentageof each component cost in relation to overall cost as production reaches a steady state.

Figure 29 - Production Cost and Revenue vs. Number of Aircraft Produced

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Figure 30 - Cost Per Unit vs. Number of Aircraft Produced

Figure 31 - Component Cost as Percentage of Overall Steady-State Cost

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Concept Comparison

Table 24 - Concept Comparison

The Socata TBM 7008 and Pilatus PC 129 are both single engine turboprops that are partof the target market. It is important that the concept has similar capabilities and costs,otherwise the benefit of using an alternative fuel will be outweighed by poor aircraft performance or a prohibitively high acquisition cost. As can be seen in Table 24 -Concept Comparison, Yamasan 2006 demonstrates similar capabilities at lower costswhen compared with its two main competitors. The design range for the concept aircraftis 800 nautical miles, including the 200 nautical mile fuel reserve. This is less than bothof the two competitors, which have a range of about 1000-1200 nautical miles. This mayexplain the difference in aspect ratios between the aircraft. The shorter design range was

selected during the QFD formation and is based on information that the largest percentage of flights cover distances less than 500 nautical miles. This justified theselection of a shorter design mission range. Another difference between the conceptaircraft and the two competitors is the power loading. The concept aircraft is intended tobe cruising at a higher speed, which may account for this difference.

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Conclusion

The Yamasan 2006 design process has produced an aircraft which meets or exceeds eachdesign goal set in the beginning of the project. Additionally, the performance is

comparable to the existing competition. Given that the Yamasan 2006 is meant to bepowered by alternative fuel, an aircraft operator should have increased incentive topurchase. While the aircraft has only undergone a simple design process, it has shownpromise in becoming a successful and practical real-world aircraft based on theoretical,mathematical, and conceptual models.

The primary goal of this design project was to design a concept which could beintroduced into a selected market in this case, the general aviation market and providethe owner with a stable and profitable solution to their business needs. The air taxi andair charter companies have been specifically targeted as potential customers. Afterextensive aerodynamic, structural, and cost analysis, it has been determined that the

Yamasan 2006 would be a highly desirable aircraft in which to invest.

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References[1] Raymer P. Daniel, Aircraft Design: A Conceptual Approach 3rd Edition, AIAAInc, Virginia 1999

[2] Scott Jeff, NACA Airfoil Series, Aerospaceweb.org, 26th August 2001http://www.aerospaceweb.org/question/airfoils/q0041.shtml

[3] Abbott and von Doenhoff, Theory of Wing Sections, Dover Publication Inc,New York 1959

[4] JavaFoilhttp://www.mh-aerotools.de/airfoils/javafoil.htm

[5] Beechcraft Starship 2000A specificationshttp://www.starshipdiaries.com/specifications.html

[6] Meyers Aircraft Company, DAPCA IV Model, Oct 1999http://www.meyersaircraft.com/DAPCA IV/DAPCA IV Intro Page.html

[7] NASA Airframe Cost Model, Jan 21, 2005http://www1.jsc.nasa.gov/bu2/airframe.html

[8] Socata TBM 700 and 850 specifications

http://www.socata.eads.net/web/lang/en/1024/content/OF00000031800002/3/00/3180000

3.html

[9] Pilatus PC-12 specificationshttp://www.pilatus-aircraft.com/html/en/products/index_201.asp?

NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3

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http://www.aerospaceweb.org/question/airfoils/q0041.shtmlhttp://www.mh-aerotools.de/airfoils/javafoil.htmhttp://www.starshipdiaries.com/specifications.htmlhttp://www.meyersaircraft.com/DAPCA%20IV/DAPCA%20IV%20Intro%20Page.htmlhttp://www1.jsc.nasa.gov/bu2/airframe.htmlhttp://www.socata.eads.net/web/lang/en/1024/content/OF00000031800002/3/00/31800003.htmlhttp://www.socata.eads.net/web/lang/en/1024/content/OF00000031800002/3/00/31800003.htmlhttp://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3http://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3http://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3http://www.aerospaceweb.org/question/airfoils/q0041.shtmlhttp://www.mh-aerotools.de/airfoils/javafoil.htmhttp://www.starshipdiaries.com/specifications.htmlhttp://www.meyersaircraft.com/DAPCA%20IV/DAPCA%20IV%20Intro%20Page.htmlhttp://www1.jsc.nasa.gov/bu2/airframe.htmlhttp://www.socata.eads.net/web/lang/en/1024/content/OF00000031800002/3/00/31800003.htmlhttp://www.socata.eads.net/web/lang/en/1024/content/OF00000031800002/3/00/31800003.htmlhttp://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3http://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3http://www.pilatus-aircraft.com/html/en/products/index_201.asp?NavL1ID=31&NavL2ID=194&NavL3ID=200&NavL4ID=0&NavL5ID=0&NavL6ID=0&L=3


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Merit Pool

Team Member Merit

John Collins 14.5Chad Davis 7.5

Chris Fles 14

Danny Sze Ling Lim 14.5

Justin Rohde 12.5

Ryan Schulz 7.5

Ronald Wong 14.5

Yusaku Yamashita 15

Total 100

Documents

Group 6 PDR Document