Group 5 SDR Document

8/3/2019 Group 5 SDR Document

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System Definition Review

AAE 451 Aircraft DesignTeam V

Stephen Beirne Charlie RushMiles Hatem Zheng Wang

Chris Kester Brandon Wedde

Jim Radtke Greg Wilson


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Team V; 2

Executive Summary

The initiative to design an alternative fuel aircraft has been undertaken to decrease thedependence of the general aviation market on fossil fuels. Team V has chosen to design a single-engine, four-seat aircraft as a lucrative inroad into this market under the confidence that peak

oil will occur in the next century, and the phase-out of 100 octane low-lead (100LL) aviationgasoline in the near future. While the market is forced to adapt to these events, this teamssolution will have already been proven as a successful product for the general aviation public.Careful attention to customer attributes in the earlier stages of design have lead to designrequirements that are certain to transcend the anticipated events and serve the general aviationmarket well into the future. With the design requirements defined, a parallel effort wasundertaken to model both the cost and size of such an aircraft.

To assure the most accurate prediction of cost, three cost estimating relationships were perfectedand employed to predict an acquisition cost of $316,000. A cost model was created based uponGTOW and max speed of a well defined aircraft database. The DAPCA and AMPR models

were modified to apply to general aviation and used to further increase the accuracy of theprojected cost. With more confidence in the teams cost modeling approach, sizing wasundertaken knowing the full implications of all design parameters.

Proper sizing was assured by using only the most educated of assumptions and investigatingdesign alternatives such as the use of composites and more radical aerodynamic shaping. Basedupon the teams customer attributes, the design mission was used as the input for the sizingmodel and all weight fractions were determined. Further detail was added to the sizing modelwith the creation of a constraint diagram to determine the possible design space and optimalpower and wing loadings. With the initial sizing of the aircraft completed, a concept had to bechosen before any detailed sizing began.

Concept selection occurred through the use of two iterations of Pughs method. All teammembers participated in this process to refine all suggested concepts to one final selection. Theconcepts ranged from an initial datum resembling a Cessna 172 to radical designs such as a bio-diesel box wing and a dual tail electric hybrid. The final iteration of Pughs method produced abase concept with a few options requiring more detailed study. This concept, outlined below,will be carried forward to the Preliminary Design Review and refined into a final systemdefinition.

Team V System Definition Specification

Power plant 200 hp tractor style bio-diesel engineWing 160 ft2 low-mounted wing with or without winglets

Tail Conventional style

Landing gear Tricycle layout w/ aerodynamic strut and fairings

Construction Aluminum frame w/ aluminum or composite skin

Cost $316,000


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Team V; 3

Table of Contents


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Team V; 4

Introduction

This teams system requirements review described a plan to target the general aviation marketwith a single-engine, four-seat aircraft was improved upon for the system definition. It is thegoal of Team V to design a product that will be marketable to hobbyists, fixed base operators,

and training fleets in the phase-out of 100 octane low-lead (100LL) aviation gasoline and thetransitional times of peak-oil. The following sections outline the method used to determine thesystem definition. A comprehensive cost model was first determined to help the team design tothe customer attribute of affordability. Preliminary sizing was started by determining thenecessary weight fractions to complete the design mission, and then followed by additionalsizing using a constraint diagram to determine the optimal power and wing loading. Throughoutthe entire sizing process, assumptions had to be carefully managed and researched to keep themodel as accurate as possible. Using Pughs method a concept was chosen as the final systemdefinition to be completely designed to accomplish the design mission.


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Team V; 5

Design Requirements

The current design requirements shown in Table 1 are still the same as presented in this teamsSystem Requirements Review. Through the work presented in this report, it will be shown that atthe current level of detail an alternate-fueled aircraft can meet all the given requirements.

Table 1 - Current Design Requirements

Design Requirements Summary

1500 ft takeoff distance to clear a 50 ft obstacle

600 lb payload with max fuel

125 kts max cruise speed, Target = 150 kts

500 nm range, Target = 600 nm

48x44 cabin height x width

Target GTOW of 2800 lbs

Target base price of $300,000


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Team V; 6

Modeling Techniques Sizing

Initial sizing was done through the use of a weight fraction method implemented in a Matlabscript. Gross takeoff weight (GTOW) was calculated using the following equation in which W0represents the GTOW, Wcrew and Wpayloadare the load weights, Wf/W0 is the fuel weight fraction

and We/W0 is the empty weight fraction:

+=

00

0

1W

W

W

W

WWW

ef

payloadcrew

(1)

To provide a reasonable estimate of the empty weight fraction, data from existing aircraft wereused to develop a predictive model. From the teams library of general aviation (GA) aircraft, asubset of 13 was selected on a basis of similarity to the teams design criteria and sales of eachaircraft. The selected aircraft and some relevant data are given in Table 2.

Table 2: Basis for empty weight fraction approximation, in order of increasing We/W0

Aircraft Range [nm] Cruise [kts] Empty Weight [lbs] GTOW [lbs] We/W0Piper Warrior III 513 896 1544 2440 0.633

Cessna 182s 821 1130 1970 3100 0.635

Diamond DA-40 583 908 1627 2535 0.642

Cessna 172S 518 914 1644 2558 0.643

Columbia 300 1130 1200 2200 3400 0.647

Piper Archer III 523 873 1677 2550 0.658

Cirrus SR-22 688 182 2250 3400 0.662

Mooney Ovation 2 1565 1108 2260 3368 0.671Columbia 350 1130 1100 2300 3400 0.676

Mooney Ovation 913 1065 2303 3368 0.684

Cirrus SR-20 720 930 2070 3000 0.690

Mooney Bravo 783 1030 2338 3368 0.694

Columbia 400 1130 1100 2500 3600 0.694

Among these aircraft, a strong correlation was found with cruise speed and empty weightfraction. The empty weight fraction also correlated with range and GTOW. Thus, a least squaresmethod was employed to build an approximation function based on the cruise speed, range,

GTOW, and empty weight fractions of the 13 selected aircraft. The resulting function is given byEquation 2, with range in nautical miles, cruise speed in knots, and GTOW in pounds:

02646.016843.0009826.0

0

][][][37328.0 = GTOWSpeedCruiseRange

W

We(2)

To verify the validity of this model, the predicted empty weight fractions for the 13 similaraircraft are plotted against the actual empty weight fractions for these aircraft in Figure 1. From


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Team V; 7

the figure, the model predicts the empty weight fraction with qualitative accuracy for bothaluminum and composite airframes. However, for the broader range of GA aircraft in the teamsentire database, the model gives less accurate results. These predictions, compared with theactual empty weight fractions, are shown in Figure 2.

Accuracy of We/Wo prediction

0.65

0.66

0.67

0.68

0.69

0.7

Actual We/Wo

PredictedWe/W

o

Alu

Ide

Co

Figure 1: Verification of empty weight fraction model


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Team V; 8

0.7

Figure 2: Verification of empty weight fraction model using entire database

The fuel fraction 0W

Wf was calculated using the equation:

=

11

2

0

1

0

...101.1x

xf

W

W

W

W

W

W

W

W(3)

Where0

1

W

Wetc. are weight fractions for the teams design mission. These fractions are either

estimated using numbers from Raymer [] or with the Breguet range equation:

= DragLift

Speed

SFCRange

i

i eW

W

1

(4)

and/or the Endurance equation:

= DragLift

SFCEndurance

i

i eW

W

1

(5)

The Wcrewand Wpayloadwere taken from the teams design requirement of having a 600 lb payloadincluding crew.


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Team V; 9

Other inputs to the code are summarized in Table 3. For the complete Matlab script, seeAppendix A.

Table 3 - Sizing Code Inputs

Cruise Speed [kts] 150Range [nmi] 600Lift-to-Drag Ratio 12Endurance Time [hr] 0.75Propeller Efficiency @ Cruise 0.86BSFC [hr-1] 0.42

Note that BSFC is the brake horsepower specific fuel consumption which is then converted intoa standard SFC using the following equation:

p

SpeedBSFCSFC

=550

(6)

where p is the propeller efficiency and the speed is in units of feet per second. Thedetermination of the BSFC is described later in this report.

The sizing code was also programmed such that if one of the inputs was made to vary, tradeoffstudy plots would be generated that would compare the varied input and both GTOW andacquisition price. Some example tradeoff study plots follow:

0.15 0.2 0.25 0.3 0.352300

2400

2500

2600

2700

2800

SFC

W0[lbs]

0.15 0.2 0.25 0.3 0.35210

220

230

240

250

260

270

280

290

SFC

TrendPrice[$k]

Figure 3 - SFC Tradeoff Study


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Team V; 10

0.6 0.8 12700

2800

2900

3000

3100

3200

3300

3400

Prop Efficiency

W0[lbs]

0.6 0.8 1250

300

350

400

450

Prop Efficiency

TrendPrice[$k]

Figure 4 - Prop Efficiency Tradeoff Study

200 400 600 800 10001000

1500

2000

2500

3000

3500

4000

4500

5000

Payload Weight [lbs]

W0[lbs

]

200 400 600 800 10000

200

400

600

800

1000

Payload Weight [lbs]

Tren

dPrice

[$k]

Figure 5 - Payload Weight Tradeoff Study (weight includes pilot)

Comparing Figure 3 Figure 4 to Figure 5, one can notice that the payload weight has much moreimpact on the GTOW and price than the SFC or the prop efficiency across the ranges shown.These ranges are intended to span the spectrum of feasibility for each variable. Fortunately,however, the original design came very close to meeting the teams design goals and not many

tradeoffs were necessary.

For completeness, Figure 6, Figure 7, and Figure 8 show tradeoff studies of the other parametersvaried through their respective reasonable ranges.


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Team V; 11

8 10 12 14 162600

2700

2800

2900

3000

3100

3200

3300

3400

3500

Lift/Drag

W0[lbs]

8 10 12 14 16250

300

350

400

450

Lift/Drag

TrendPrice

[$k]

Figure 6 - Lift-to-Drag Ratio Tradeoff Study

100 120 140 160 180 2002200

2400

2600

2800

3000

3200

3400

Cruise Speed [kts ]

W0[lbs]

100 120 140 160 180 200200

250

300

350

400

450

Cruise Speed [kts ]

TrendPrice[$k]

Figure 7 - Cruise Speed Tradeoff Study

200 400 600 800 10002200

2300

2400

2500

2600

2700

2800

2900

Range [nmi]

W0[lbs]

200 400 600 800 1000180

200

220

240

260

280

300

Range [nmi]

TrendPrice[$k]

Figure 8 - Range Tradeoff Study


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Team V; 12

The final run of this sizing code yielded a GTOW of 2834 lbs. This includes a calculated fuelweight of 358 lbs and an empty weight of 1865 lbs.

To verify the accuracy of the sizing methodology, the input parameters were changed to reflect

two of the benchmark aircraft, the Cessna 172 and Cirrus SR22. The inputs and results arepresented in Table 4. These results validate the sizing code, showing that it provides reasonablepredictions of aircraft weight.

Table 4: Sizing code validation

Aircaft Cessna 172 Cirrus SR22

Input Parameter

Cruise Speed [kts] 122 185

Range [nmi] 580 700

Lift-to-Drag Ratio 10 14

Endurance Time [hr] 0.75 0.75Propeller Efficiency @ Cruise 0.86 0.86

BSFC [hr-1] 0.45 0.4

Full Fuel Payload [lbs] 482 664

Results of Sizing Code

Output Weight [lbs] 2246 3385

Actual GTOW [lbs](From Manufacturer) 2450 3400

% Error 8.3% 0.4%


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Team V; 13

Modeling Techniques - Acquisition Cost Estimation

A very important aspect of this project is being able to predict and maintain the quality of aprediction of acquisition cost. One of the best ways to predict cost at the level of this project isby means of cost estimating relationships (CER). There were three acquisition CERs employed

in determining the current predicted cost of $316k

The first model utilized is one based upon the general aviation library created by the team forthis project. The library contains a variety of single-engine piston-prop-powered aircraft from theLiberty XL2 to the Piper Saratoga 2TC. It is understood that this is quite a range of aircraft, withacquisition costs ranging from $140k to $570k (2006 $), however it was reasoned that such avariety of aircraft could reinforce the models created, barring any unforeseen trends that did notmake intuitive sense.

A few relationships discovered with this CER are presented in Figure 9 and Figure 10below.The associations represented are GTOW and max speed vs. acquisition cost. In each case, the

purchase price is presented as is, along with an exponential trend line.

Figure 9 - Acquisition Cost v. GTOW


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Team V; 14

$700 000.0

)

Figure 10 - Acquisition Cost v. Max Speed

The findings from the GA library seem to make intuitive sense: as takeoff weight and max speedincrease, acquisition cost follows suit.

As mentioned before, two other models were used. One used was a version of the RANDDAPCA model, as presented in chapter eighteen of Raymer []. The model traditionally takes theinput presentedbelow in Table 5. Table 5 also contains the constant values affixed to theirrespective parameters. Since it is not a turbine aircraft but rather a piston aircraft, the model wasmodified such that the engine production cost was not factored into the CER. This is areasonable assumption in the fact that the end product, the output of purchase price, needed to bemultiplied by a factor to relate this CER designed for military aircraft to aircraft of the scope ofthis project. If engine production cost had been figured into the model, the multiplication factorwould merely be larger. This multiplication factor also accounts for the update of the model to2006 dollars in the same manner.


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Team V; 15

Table 5 - DAPCA CER Input

Weight GTOW --

Empty weight --

Speed Max velocity --

Production Total Quantity 5000

5 year Goal 2500

Number of flight test aircraft 1

Turbine Engine Number of engines --

Max thrust --

Max mach --

Turbine inlet temp --

Avionics* Avionics mass 5 kg

Each aircraft of the teams database was run through the DAPCA model with an investmentfactor of 1.1 (10% profit per unit). A few examples of output are presented in Table 6 below.

Table 6 Sample DAPCA Output

Aircraft Acq. Cost DAPCA Prediction

Cessna 172S $ 241,000.00 $ 452,252.69

Cessna 182s $ 326,150.00 $ 563,444.45

Cirrius SR 22 $ 448,685.00 $ 717,619.85

It is readily apparent that the model was a poor predictor as is, overestimating by as much as70%. To alleviate this problem, a fudge factor was introduced. The actual price of the aircraftwas on average 59% of the DAPCA output, therefore it was clear that if the output weremultiplied by the above percentage, that it might better represent the Team V aircraftsclassification. Once done, the average agreement between prices was indeed an even 100%. Theresulting data is displayedbelow in Figure 11 where it can be seen that the DAPCA model asmodified above is a reasonable estimator of cost within the target GTOW range (2500 - 3000lbs).

* Avionics cost was added into the DAPCA model as suggested in the text. Avionics account for ~5% of flyawaycost at ~$4500/lb. []


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Team V; 16

$1,000,00Figure 11 - Purchase Price CER v. DAPCA CERs

Accounting for possible increased costs due to composite construction was also a goal of themodified DAPCA model; however, it was found that there was no change in the curve fitthroughout the entire range of 1.1-1.5 multiplication factor to account for the increased costs.

In addition to the DAPCA model, a third model using estimates of airframe unit weight cost.This CER appears again in chapter 18 of Raymer [] and was optimized for the presentapplication. According to the text, the 1999 $ cost-per-pound of airframe unit weight (AMPRweight) was between $200 and $400. This AMPR weight is approximately 60 70% of theaircraft empty weight and for this application was set at 65%.

The only aspect of the CER remaining was to pick the value between $200 and $400 that wouldbest fit the GA library. Three approaches to the optimization were taken: maximize the averagepercent agreement [Average(actual/CER) =100%]; equaling the sum of the difference to zero

[Sum(actual-CER) = 0]; and minimize the maximum difference [Max(actual-CER)]. MicrosoftExcels Solver with default settings was utilized to complete to optimizations. The valuesdetermined are presented in Table 7 below.


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Team V; 17

Table 7 AMPR CER Optimization results

Method $/AMPR

Average % Difference = 100% $ 257.25

(Price Differences) = 0 $ 277.05

Minimize (Max. Price Difference) $ 260.37

After completing the optimizations, it was determined that equaling the sum of the pricedifferences to zero yielded the best agreement with the actual purchase prices and the DAPCACER. All three final cost estimating relationships for acquisition cost are presented in Figure 12and the equations for the exponential trend lines are presented in Table 8.

Figure 12 - Finalized CERs

Table 8 Trendline and2

R values for utilized CERs

CER Exponential Trendline 2R

Purchase Price xey 0006.047622= 0.86

DAPCA xey 0004.092212= 0.86

AMPR xey 0004.0107573= 0.95


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Team V; 18

It is apparent that the AMPR CER has the best 2R value, but this is most likely because it is the

simplest relationship used, breaking the complex association between aircraft and purchase pricedown to two parameters AMPR and $/AMPR. The other CERs are still very good. It isinteresting to see that all three intersect at approximately (3250 lb; $375,000).

Future plans for costing the aircraft include completing development of a lifecycle cost model.The lifecycle cost model would include purchase price, special construction costs, operations andmaintenance costs along with disposal costs. Operations and maintenance costs would includefuel, maintenance, insurance and depreciation. The current model estimates that at a purchaseprice of $330k (2006 $), upkeep-per-year totals about $60k and the lifecycle cost of the aircraftover 12 years would be approximately $950k. It would be the goal of this team to aim for apurchase price and lifecycle cost that would allow for the main target commercial customers(fixed based operators and training fleets) to make at least a 10% return on their purchase andinvestment per year ($390k for the first year yields the customer $60k + $39k). This wouldensure that purchase of this teams aircraft would be a better investment than investing moneyelsewhere. For the hobbyist market, the goal would be to maintain a 10% return on fun and

enjoyment.


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Team V; 19

Modeling - Fuel Consumption Modeling and Selection

An important part of the engine design of the aircraft was to find brake horsepower specific fuelconsumption for the fuel and engine the group chose. To determine this choice, horsepowerspecific fuel consumptions were calculated and used as a selection criterion. To do this a table

was constructed of the different types of possible fuel sources and the characteristics of each ofthese fuels. This is shown in Table 9.

Table 9 - Characterized Possible Fuel Sources

Fuel Specfic Gravity Density [lb/gal] Combustion Energy [Btu/lb]StoichiometricAir-fuel ratio

Octane 0.7 5.87 20750 9Diesel 0.85 7.09 18300 15Ethanol 0.79 6.58 13160 15.2Biodiesel 0.88 7.34 16000 13.8Jet A 6.46 18700Hydrogen -- 61000

The team also previously decided to use a piston engine. A diesel engine was initially chosen asa baseline because it has lower specific fuel consumptions than both turbine and typical gasolinepowered engines, is less complicated than the other two engines due to the nature of the dieselcycle, and it produces more power at lower RPMs which make it suitable for a propeller drivenaircraft. A database of existing aircraft diesel engines was created for a starting point incalculating the characteristics of diesel engines running on alternative fuel. The database ofdiesel engines is shown in Table 10.

Table 10 - Diesel Engine Database

Name Power [hp] BSFC [lbs/hp*hour] Weight [lbs] Engine Volume [ft 3]Thielert Centurion 1.7 135 0.36 295.41 8.31Thielert Centurion 4.0 310 0.36 606.3 39.16DeltaHawk V-4 Diesel 160 0.4 327 12.11DeltaHawk V-4 Diesel 200 0.38 350 15.39Zoche ZO 01A 150 0.365 185 5.86Zoche ZO 02A 300 0.355 271 16.70Zoche ZO 03A 70 0.381 121 1.87

No data for break specific fuel consumptions (BSFC) for alternative fuels in diesel engines wereavailable at the time of this report. To determine the BSFC for each alternative fuel, an average

of the BSFCs were taken from the diesel engine library. This information was used along withthe heating value of number two diesel fuel and its density to find a power requirement perhorsepower per hour. This number was then used with the heating values and densities of thealternative fuels to find an estimated BSFC for each alternative fuel. The following samplecalculation demonstrates this process for finding the BSFC for Bio-Diesel:


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Team V; 20

Determining the power requirement for a diesel engine:

hourhp

Btu

gal

Btu

lbs

gal

hourhp

lbs

*8508.6848129500*

0936.7

1*

*37516. =

BSFC Calculation for 100% Bio-Diesel:

hourhp

lbs

gal

lbs

Btu

gal

hourhp

Btu

*425.344.7*

118296

1*

*8508.6848 =

These and similar calculations resulted in the values shown in Table 11 for each of thealternative fuels analyzed.

Table 11 - Fuel BSFC Comparison

Fuel BSFC (lbs/hp*hour)

Diesel #2 0.37516Bio-Diesel B100 0.42518

Ethanol 0.53829Liquid Hydrogen 0.1269

From this it was found that hydrogen gave the best BSFC, but hydrogen was ruled out because ofthe goal to replace 100LL aviation gasoline. Since 100LL will be phased out in about ten years,hydrogen technology will not realistically be advanced enough to use hydrogen as a fuel.Hydrogen will require the addition of large tanks and pumps that would not be a good fit on asmaller general aviation aircraft. The addition of this extra equipment will cause a large increasein GTOW for an airplane, which in effect would most likely triple or quadruple the estimatedBSFC for hydrogen. That would make this fuel choice less competitive with the otherpossibilities. Hydrogen also has no existing distribution infrastructure, which would only further

delay its implementation as a readily available fuel due to its unique storage and transportationrequirements.

The next best BSFC was bio-diesel, which was selected as the alternative fuel for this aircraft.The team chose Bio-Diesel for several reasons. First, its BSFC is only slightly lower thanregular diesel fuel. Second, using Bio-Diesel requires only minor modifications to the engine itwill fuel. At times this fuel can soften rubber hoses, but this is easily fixed by replacing thosehoses and is only a problem in older engines. The fuel can initially clog fuel filters because it

Average BSFCfrom engine library

Density of #2Diesel

Heating Value of #2Diesel

Average Power Requirementfor a Diesel Engine

Heating Value of

Bio-Diesel

Density of Bio-Diesel


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Team V; 21

has a solvent action that will dissolve deposits in an older vehicles fuel tanks. This is easilyremedied by replacing the filters, however the engine will be a new product designed to run onBio-Diesel, so this aircraft will not face any of these problems. Third, the supply of Bio-Dieselis readily available and the amount refined is growing rapidly each year. Between 2004 and 2005the quantity of Bio-Diesel refined grew by a factor of three. Fourth, Bio-Diesel does not have

any special handling requirements like hydrogen or ethanol. This makes it easy to transport andmuch of the already existent infrastructure can be used to distribute it in the future. Last,although Bio-Diesel does experience cold flow problems like regular diesel, this can be solvedby the use of additives to lower the gel point to a desired level, or a tank heater can be installed.

Bio-Diesel was chosen over ethanol because ethanol had a higher specific fuel consumption.The fuel is also corrosive to any rubber hoses on an engine and can require more modification toan engine than that for Bio-Diesel.

Further, estimates of engine sizing were prepared to facilitate future work on this aircraft. To getan idea of the weight that the engine would add to the aircraft, a plot was made using the

information from the diesel engine database. From this plot a linear relationship was foundbetween the horsepower of the engine and the overall weight of the engine. As expected, as thehorsepower increases so does the weight of the engine. This plot is shown in Figure 13.

Power vs. Weighty = 0.4563x + 27.501

R2= 0.912

0

50

100

150

200

250

300

350

100 150 200 250 300 350 400 450 500 550 600 650

Weight (lbs)

Power(h

Figure 13 - Engine Horsepower vs. Weight

For the next part of this project, this group would like to get a volumetric sizing of the engineand, given a power requirement, predict the weight that it will add to the aircraft. The groupwould also like to determine the number of cylinders and possibly the volume of the cylindersneeded to output the required horsepower.


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Team V; 22

Constraint Diagrams

To find the wing loading, engine power, and wing dimensions of the concept aircraft, aconstraint diagram was created. Six constraints including take-off, cruise, landing, climb, turn,and stall speed were used to determine a design space from which a design point could be

chosen. Using definitions, algebra, and several equations from Raymer [1], the power-to-weightratio, as a function of wing loading, was found for each of the six constraints. A Matlab scriptwhich included the six power-to-weight ratio equations for the constraints was written and eachequation was run for a range of wing loadings. For each value of wing loading, a value ofpower-to-weight ratio was calculated and the values for each constraint were plotted on the sameaxes. Figure 14 below shows the completed constraint diagram.

0 5 10 15 20 25 30 35 400

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.1

Wing Loading, WTO

/S

Power

toWeightRatio,

PSL

/W

TO

Constraint Diagram

Take-Off

Cruise

Landing

Climb

Turn

Stall

Figure 14 - Constraint Diagram

Curves were created for each constraint by plotting the power-to-weight and wing loadingvalues. These curves intersect with one another to define a design space that satisfies eachconstraint. For cruise, climb, take-off, and turning constraints, all points above each line satisfythe constraints. For the landing and stall speed constraints, all points to the left of eachconstraint curve are solutions. With the design space determined, a design point can be chosenwhich will optimize the power-to-weight ratio and wing loading for the aircraft. As can be seenfrom Figure 14, the design point was chosen to be at the lower corner of the design space. At the

Design Point

DesignSpace


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Team V; 23

design point, a wing loading of 17.8 lb/ft2 and a power-to-weight ratio of 0.07 satisfy all sixconstraints.

It should be noted that in the creation of the constraint diagram many modeling assumptionswere made to simplify the design process. Table 12 below lists the input parameters and their

values. Several of the assumed parameters were varied to allow the power requirement to remainbelow 200 hp for the aircraft. At any higher horsepower the aircraft would be considered highperformance, preventing its use in the training market.

Table 12 - Constraint Diagram Parameters

Parameter Value

Stall Speed [kts] 57

Cruise Speed [kts] 150

CLMax 1.6

Take-off Roll [ft] 740

Prop Eff. @ Take-off 0.66

Prop Eff. @ Cruise 0.86

Landing Roll [ft] 740Aspect Ratio 8.6

Parasite Drag 0.026

Climb Rate [ft/min] 1000

Turn Load Factor 2

With the estimated max takeoff weight of 2834 lbs determined from the sizing code that waspreviously mentioned, the constraint diagram was scaled to show horsepower and wing area.Figure 15 shows the scaled constraint diagram, in which the design space is in the upper-rightcorner.


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Team V; 24

100 110 120 130 140 150 160 170 180 190 200100

120

140

160

180

200

220

240

260

280

300

Wing Area [ft2]

HorsePower

Scaled Constraint Diagram

Take-Off

Cruise

Landing

Climb

Turn

Stall

Figure 15 - Edited constraint Diagram

At the design point, the designed aircraft would require an engine with about 200 horsepowerand a wing area of around 160 ft2. Using the input aspect ratio and the calculated wing area, thewing span of the aircraft was determined to be about 36 ft.

DesignSpace

Design Point


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Team V; 25

Pugh's Method

Pughs method is a fast and effective way to iterate through a large number of initial concepts, toeliminate inferior designs, and to promote superior designs. Thus, it was employed to develop aconcept for the aircraft for this project.

Each team member produced an initial concept to start Pughs method. The concepts variedwidely, and included such features as canards, conventional tails, T-tails, low wings, high wings,glass cockpits, solar-cell wing skins, pusher props, fixed gear, retractable gear, and many more.Additionally, two concepts based on benchmark aircraft, the Cessna 172 and Cirrus SR22, wereincluded.

A table was constructed such that each concept was evaluated against a datum aircraft based on anumber of criteria, based on the customer and engineering attributes from the SystemRequirements Review. The criteria included:

Engine Power GTOW

Cruise Speed

Safety

Airport Compatibility

Price

Payload

Pizzazz

Stability

Fuel Consumption

Aesthetics

After each iteration of the method, the clearly bad concepts were eliminated, and good conceptswere merged. The process was then repeated until the number of concepts was reduced to two.The two final concepts are described in the Selected Concepts section.

Both final concepts are variations on the same basic concept, sharing all characteristics exceptwinglets, propeller type (fixed vs. variable pitch), and construction material.

The tradeoffs involved in winglets were discussed extensively, and it was determined thatdetailed analysis was needed to finally select a concept. Though winglets may reduce drag forlarge, fast aircraft, the slow and small aircraft under consideration here would not reap muchaerodynamic benefit. The potential benefit of winglets in this case comes in the realm ofmarketing. Winglets are visually appealing and tend to help aircraft sell, even when theperformance benefits are not compelling. The disadvantages of winglets are the weight and costpenalties incurred to construct the more complex wing.

The choice between fixed and variable pitch propellers was difficult as well. Variable pitchpropellers provide better performance, but at a cost, weight, and complexity disadvantage to


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fixed-pitch propellers. Thus, it was decided to allow both concepts to remain, until more detailedanalysis could better differentiate them.

Three different concepts for the construction of the aircraft were included within the variousconcept aircraft evaluated with Pughs Method: all aluminum construction, all composite

construction, and a hybrid concept with composite skin bonded to an aluminum frame.

Low weight and low cost were expected to be the deciding criteria for the structural concepts. Itwas also thought that different structural choices may confer a performance difference.

To truly evaluate the concepts on the basis of cost, detailed structural analysis and processevaluation beyond the scope of this phase of the project are required. However, initial judgmenton the cost of producing each concept allowed the all-composite variant to be eliminated. It isthought that the thicker components of the internal structure, such as the spars, ribs, and cabinframe, could be more easily fabricated with aluminum than composites. The lay-up of carbon orglass fiber prepregs into the thickness that would be required in a spar would require 30 or more

plies. Further, the typical I-beam spar cross section is difficult to achieve with conventionalpractices. For a general aviation aircraft, the added cost of such build-up of thick parts withcomposites makes aluminum a more attractive material for the internal structure.

Thus left with the aluminum and hybrid structural concepts, weight and performance wereexpected to provide a distinction between them. However, an examination of the aircraft in theteams database, comparing the aircraft on a basis of material composition, showed that neitheraluminum nor composite aircraft were inherently superior in either aspect of weight orperformance.

For example, Figure 16 shows the empty weight fraction of every aircraft in the database,grouped by material. There are low and high empty weight fractions in each group, and none hasa clear edge.


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Figure 16: Empty weight fraction by construction material

It is possible that the empty weight fraction does not tell the full story, since structural weightsavings may be used to add systems or engine weight. Figure 17 shows the empty weightfraction plotted against engine horsepower. Since more powerful engines generally weigh more,the trend shown is that of increasing empty weight fraction with horsepower. Aircraft with betterstructural weight have lower empty weight fractions despite larger engines. From the figure, it isevident that aluminum aircraft have a slight advantage in this respect, but again both materials

offer good and poor examples.

It should be noted that the hybrid aircraft shown are the Liberty XL2 and Symphony 160, both ofwhich are low-end two-seat aircraft. Thus, their poor performance in these comparisons shouldnot be construed as weakness of the hybrid concept overall. To date, no aluminum/compositehybrid structure aircraft in a more applicable class are available for comparison.

Composite Aluminum Hybrid


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Engine Power Effect on We/Wo

0.4

0.45

0.5

0.55

0.6

0.65

0.7

0.75

0 50 100 150 200 250 300 350 400

Engine HP

We/Wo Metal

Composite

BothBetter

Figure 17: Power tradeoff with empty weight fraction.

As a measure of performance, in Figure 18 the acquisition cost is shown versus a representationof performance, the product of cruise speed and range. Here again, there is no definitivelysuperior material; however, the composite aircraft appear to have a slight advantage.


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Cost of Performance

$0

$200,000

$400,000

$600,000

$800,000

$1,000,000

$1,200,000

0 50000 100000 150000 200000 250000 300000 350000

Range*Speed [nmi^2/hr]

Price[$]

Metal

Composite

Both

Figure 18: Bang-for-the-buck comparison of price paid for the performance measure combining range and speed

Thus, with no clear differentiating factor based on historical trends, both the all aluminum andthe hybrid concepts will be carried forward, until detailed design work and cost analysis allowfor a single concept to be selected.


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Selected Concepts

After two iterations of Pughs method, two very similar designs emerged as the winners. Theattributes of each design are listed in Table 13.

Table 13 - Selected designs after two Pughs iterations

Concept A Concept B

Landing Gear

Fixed Gear/FaringTricycle Gear

Fixed Gear/FaringTricycle Gear

Wing

Low WingStraight Wing

Low Wing with WingletsStraight Wing

Tail Conventional Tail Conventional Tail

Cockpit Avionics Analog Cockpit Glass Cockpit

Construction Al Al frame composite skin

Propeller Fixed Pitch Variable Pitch

Fuel Bio Diesel Bio Diesel

Concept B was chosen as an upgraded version of Concept A, with better performance but highercost. In both concepts, fixed tricycle gears were chosen for their reliability, safety, lower weight,and cost compared to retractable gears. It was determined that, at the required cruise speed, theadditional drag is low enough to be acceptable. Low, straight wing configuration was chosen forits structural advantages. Low wing allows the wing spar to be directly integrated into thefuselage, without the need for supporting braces typically required by high wing configurations.A low wing also allows the landing gear to be mounted on the strong wing box structure, whichshortens the landing gear and eliminates need for extra structures in the fuselage, thus reducingweight. The conventional tail configuration was chosen for its simplicity, safety and lower

weight compared to more elaborate designs. Both designs use Bio-Diesel, since the fuel analysisshowed that it is the best choice.


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Preliminary Fuselage Layout

Figure 19 - Preliminary Cabin Layout

The cabin seating layout is the classic two-by-two layout seen in most small GA aircraft. Theseating arrangement is designed to satisfy the 46 x 48 interior dimension, and is produced withthe help of Catia V5s human activity analysis work bench. Figure 19 shows a three view of thecabin, and its relation with respect to the airplane exterior.


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3-View of Best Current Design

Figure 20 - Three View of Current Aircraft Design


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List of current aircraft parameters and performance projections

Engine:

200 Hp Bio Diesel

Dimensions:

Wing span: 36 ftLength: 29 ftHeight: 9 ft

Wing area: 160 ft2

Aspect ratio: 8.6

Performance

Range: 600 nmCruise speed: 150 ktsTakeoff distance (clear 50 ft obstacle): 1500 ftTakeoff ground roll: 740 ft

Weight

Max takeoff: 2834 lbPayload with full fuel: 600 lbEmpty Weight: 1865 lbFuel Weight: 358 lb

Price

Acquisition cost: $316,000 (2006 dollars)


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Comparison

To see how the designed aircraft fairs against similar aircraft in the current market, its parameterswere compared to existing aircraft. Table 14 below compares the characteristics of the Team Vaircraft to several benchmark aircraft, with which it may compete when it first enters the market.

As can be seen from the table, the aircraft defined in this paper satisfies the set design goals forall but the gross take-off weight and acquisition cost parameters. Even though these goals arenot met, they are only overshot by small margins.

A comparison of the Team V solution to that of its competitors shows that it certainly hasdefined its place in the market. Even though it has taken a slight sacrifice in range and payloaddue to the switch to an alternative fuel, it is still close to the Cessna 172. It was decided that thisrange and payload was acceptable for the target market, in which the primary customer will betraining operations and hobbyists who do not demand a longer range and large payload forutility.

Concerning performance, the aircaft is comparable to the benchmarks, with a climb rate andcruise speed similar to the Cessna 182, 182T, and Diamond DA-40. Prior Quality FunctionDeployment studies showed that an upper range cruise speed would be more attractive topotential customers. Also the aircraft has a shorter ground roll at take off which can be attributedto its higher power and wing loading compared to the lower performance Cessna 172.

Designed for both the trainer and hobbyist markets, this aircraft has been designed to have manyperformance advantages that raise it above the entry level competition at a purchase price lessthan the high performance aircraft. The Team V aircraft has remained safe while offering theentry level flyer a product that will still be exhilarating during the events of peak oil and 100low-lead phase out.


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Table 14: Comparison to competitors

Cessn

a172

Cessn

a172SP

Cessn

a182

Cessn

a182T

Diam

ond

DA

-40

Cirrus

SR

-22

Team

VAir

craft

Range @ 45% power (nm) - 638 - 971 - - -

Range @ 55% power (nm) - - 930 - - - -

Range @ 60% power (nm) 687 - - - - - -

Range @ 75% power (nm) - 518 - - 600 700 600

Range @ 80% power (nm) 580 - 773 615 - - -

Range @ 88% power (nm) - - - - - - -

Power (bhp) 160 180 230 235 180 310 200

Power loading (lbs/hp) 15.3 14.2 13.5 13.2 14.1 10.9 14.2

Climb at sea level (fpm) 720 730 924 1040 1070 1304 1000

Cruise speed (kts) 122 124 145 159 147 185 150

Stall speed, flap up (kts) 47 48 49 49 49 61 57

Ground roll (ft) 945 960 795 775 720 1020 740

GTOW (lbs) 2457 2558 3100 3112 2535 3400 2834

Usefull load (lbs) 818 895 1140 1037 906 1150 958

Wing area (sf) 174 174 174 174 145 145 160

Wing loading (lbs/sf) 14.1 14.7 17.8 17.8 14.5 23.4 17.7

Purchase Price ($) $172k $180k $326k $355k $259k $350k $316k


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Conclusion

This System Definition Review has demonstrated that, at the current level of detail, Team V hasdeveloped a concept capable of meeting all design requirements established in the SystemRequirements Review. The result of thorough sizing and cost analysis is an aircraft that

compares well with current aircraft in the general aviation market and is expected to becompetitive in the assumed future market conditions.

The final concept is a Bio-Diesel fueled, four seat, low-wing, single engine, fixed gear aircraft.Detailed concept selections, such as construction material and propeller complexity, remain forfurther work.

Team V is confident that the system definition outlined will be a high quality alternative fuelreplacement for the general aviation market. Careful attention to the customers needs, and wellthought out design will assure a solution that will perform to its design mission.

While only the foundation has been laid for the design of this aircraft, Team V is certain that itwill lead to a fully developed design capable of serving the general aviation market well into thefuture.


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References

1. Raymer, D.Aircraft Design: A Conceptual Approach. AIAA Education Series. ThirdEdition. Reston, VA. 1999.

2. Cessna SkyhawkSP. Your Next Wingtips. Cessna Aircraft Company Website. Cited 29Jan. 2006. Available: .

3. Cessna Skylane. Redefining the SUV. Cessna Aircraft Company Website. Cited 29 Jan.2006. Available: .4. Cirrus Aircraft Company Website. Cited 29 Jan. 2006. Available:

.5. Diamond Aircraft Company Website. Cited 29 Jan. 2006. Available:

.6. Velocci, Anthony L., Jr. Editor.Aerospace Source Book. McGraw-Hill Companies. Vol

164. No 3. Jan 2006.7. National Bio-Diesel Board Website. Cited 12 Mar. 2006. Available:

.


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Appendix A Sizing Code

% Initial Sizing Code% AAE 451 Spring 06% Group 5

clearclcclose all% Input sectipon

Range=600; % [nmi]AA_Range=0; % alternate airport range [nmi]Speed=150; % Cruise speed [kts]LD=12; % lift/drag ratioEndur=.75; % loiter endurance [hrs]W_crew=150; % crew weight [lbs]W_payload=450; % payload weight [lbs]

Alt=8000; % cruise altitude [ft]Prop_Eff=.86; % prop efficiencybhpSFC=.42; % brake horsepower Specific Fuel ConsumptionW0_guess=3000; % initial GTOW guess [lbs]

% loop inputstol=.0001; % convergence toleranceW0=1; % initialize W0 for loop

while(abs((W0_guess-mean(W0))/mean(W0)) > tol)% Specific fuel consumption calculation

SFC=bhpSFC.*Speed.*1.687./(550.*Prop_Eff);% Weight Ratios

w1w0=.99;w2w1=.99;w3w2=exp(-Range.*SFC./(Speed.*LD)); % range equationw4w3=1;w5w4=w1w0;w6w5=w2w1;w7w6=exp(-(AA_Range.*SFC)./(Speed.*LD)); % range equationw8w7=exp(-(Endur.*SFC)./LD); % endurance equationw9w8=w4w3;w10w9=.995;

% Weight calcs

WeW0=0.37327989.*Range.^-0.00982601.*Speed.^0.16843228.*W0_guess.^-0.0264551;WfW0=1.01.*(1-

w1w0.*w2w1.*w3w2.*w4w3.*w5w4.*w6w5.*w7w6.*w8w7.*w9w8.*w10w9);W0_guess=W0;


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W0=(W_crew+W_payload)./(1-WfW0-WeW0);% Price

TrendPrice=42127.0173690882.*exp(0.00068629368.*W0)/1000; % Estimated Sell price [$k]AMPRPrice=W0.*WeW0.*.65.*.27705;

end

% output and plotingformat longif(length(W0)==1)

fprintf('Your GTOW is %4.0f lbs \nYour database estimated price is $%3.1f k\nYour AMPRestimated price is $%3.1f k',W0,TrendPrice,AMPRPrice)elseif(length(Range)>1)

subplot(1,2,1)plot(Range,W0)

xlabel('Range [nmi]')ylabel('W0 [lbs]')subplot(1,2,2)plot(Range,TrendPrice)xlabel('Range [nmi]')ylabel('TrendPrice [$k]')

elseif(length(Speed)>1)subplot(1,2,1)plot(Speed,W0)xlabel('Cruise Speed [kts]')ylabel('W0 [lbs]')subplot(1,2,2)plot(Speed,TrendPrice)xlabel('Cruise Speed [kts]')ylabel('TrendPrice [$k]')

elseif(length(LD)>1)subplot(1,2,1)plot(LD,W0)xlabel('Lift/Drag')ylabel('W0 [lbs]')subplot(1,2,2)plot(LD,TrendPrice)xlabel('Lift/Drag')ylabel('TrendPrice [$k]')

elseif(length(Endur)>1)subplot(1,2,1)plot(Endur,W0)xlabel('Loiter Endurance [hrs]')ylabel('W0 [lbs]')subplot(1,2,2)


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plot(Endur,TrendPrice)xlabel('Loiter Endurance [hrs]')ylabel('TrendPrice [$k]')

elseif(length(W_payload)>1)subplot(1,2,1)

plot(W_payload+150,W0)xlabel('Payload Weight [lbs]')ylabel('W0 [lbs]')subplot(1,2,2)plot(W_payload+150,TrendPrice)xlabel('Payload Weight [lbs]')ylabel('TrendPrice [$k]')

elseif(length(Prop_Eff)>1)subplot(1,2,1)plot(Prop_Eff,W0)xlabel('Prop Efficiency')

ylabel('W0 [lbs]')subplot(1,2,2)plot(Prop_Eff,TrendPrice)xlabel('Prop Efficiency')ylabel('TrendPrice [$k]')

elseif(length(SFC)>1)subplot(1,2,1)plot(SFC,W0)xlabel('SFC')ylabel('W0 [lbs]')subplot(1,2,2)plot(SFC,TrendPrice)xlabel('SFC')ylabel('TrendPrice [$k]')

end


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Appendix B Constraint Diagram Code%AAE 451 Team V - Stephen Beirne, Charlie Rush, Miles Hatem, Zheng Wang% Chris Kester, Brandon Wedde, Jim Radtke, Greg Wilson%Constraint Diagramclc;close all;clear all

W_S=linspace(0,120);P_W=[0 0.5];V_stall=57*1.6878; %ft/sV_max=150*1.6878; %ft/s%Take-Off%%%%%%%%%%%%%%%%%%%%%%%%%%%%g=32.174; %ft/s^2rho=0.0024; %slug/ft^3CLmax=1.6; %Max Lift CoefficientSto=740; %fteta_p_to=0.66; %Propeller Efficiency (about 0.8)eta_p_cr=0.86;

Vto=1.1.*V_stall;T_W=(1.2./(g*rho.*CLmax.*Sto)).*W_S;Pto_W=Vto.*T_W./(550.*eta_p_to);%Landing%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%Sl=740; %ftbeta=1; %Landing weight to take-off weight ratiomu=0.4; %Coefficient of rolling friction (0.3-0.5 for dry runway)V_td=1.15.*V_stall;W_S_land=Sl.*rho.*CLmax.*g.*mu./((1.15.^2).*beta)W_S_land1=[W_S_land W_S_land];

%Cruise%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%alpha=0.75; %Percent of max power used @ cruiseV_cruise=V_max; %Velocity @ cruiseAR=8.6; %Aspect Ratio AR=B^2/Se=1; %Oswald's efficiency factorCdo=0.026; %Parasite dragrho_c = 0.001869;Psl_Wto=(beta./(alpha.*550.*eta_p_cr)).*(rho_c.*V_cruise.^3.*Cdo./(2.*beta.*W_S)+(2.*beta./(rho_c.*V_cruise.*pi.*AR.*e)).*W_S);% Climb from Raymer 5.28/5.29V_v = 1000/60;V_y = 80*1.6878;V_h = sqrt(V_y^2-V_v^2);G = V_v/V_h;q = rho*1/2*V_y^2;D_W = q*Cdo./(W_S)+W_S.*1/(q*pi*AR*e);

T_Wcl = D_W+G;P_Wcl = V_y.*T_Wcl./(550.*eta_p_to);%Turn%%%%%%%%%%%%%%%%%%%%%%%%%%%%%


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q = .5*.0024*(.8*V_max)^2;n = 2;PW_turn =.8*V_max*((q*Cdo)./W_S+W_S.*(n^2/(q*pi*AR*e)))./(550*eta_p_to);

%Stall%%%%%%%%%%%%%%%%%%%%%%%%%%%W_Sstall = .5*CLmax*rho*V_stall^2;

figure(1)plot(W_S,Pto_W,W_S,Psl_Wto,W_S_land1,P_W,W_S,P_Wcl,W_S,PW_turn,[W_SstallW_Sstall],P_W,'k')legend('Take-Off','Cruise','Landing','Climb','Turn','Stall')title('Constraint Diagram')xlabel('Wing Loading, W_T_O/S')ylabel('Power to Weight Ratio, P_S_L/ W_T_O')grid onaxis([0 40 0 .1])figure(2)weight = 2834;plot(weight./W_S,Pto_W*weight,weight./W_S,Psl_Wto*weight,weight./W_S_land1,P_W*weight,weight./W_S,P_Wcl*weight,weight./W_S,PW_turn*weight,weight./[W_SstallW_Sstall],P_W*weight,'k')legend('Take-Off','Cruise','Landing','Climb','Turn','Stall', 'location','NorthWest')title('Scaled Constraint Diagram')xlabel('Wing Area [ft^2]')ylabel('HorsePower')grid onaxis([100 200 100 300])

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