System Definition Review - Purdue Engineering 1... · Several trade studies were considered to...

System Definition Review

Team 1 Miguel Alanis Timothy Block

Becca Dale Joseph Fallon Aaron Mayne

Jason Olmstead Adeel Soyfoo Sarah Weise

March 23, 2006 AAE 451 – Senior Design

Systems Definition Review March 23, 2006

Executive Summary A rugged and versatile aircraft is currently being designed. This plane will be fueled by a

non-petroleum based alternative fuel. Possible markets for the proposed aircraft include the

European Union and Australia, who have expressed concern over emissions. Fischer-Tropsch

kerosene and biodiesel, among other fuels, have been explored as possibilities to replace Avgas

and diesel for their compatibility to current storage and power plant technology, as well as their

environment-friendly makeup. The aircraft will be powered by turboprop engines. Several trade

studies were preformed to develop initial sizing values for the aircraft. Among these were

GTOW and range trade studies. A performance constraint diagram was also created based on a

specified design mission. After analyzing the constraint diagram and comparing several existing

aircrafts’ design points, the team decided to design the plane for a power to weight ratio of 0.095

hp/lb and a wing loading of 42 lb/ft2.

A process used for concept generation and evaluation, Pugh’s Method, was applied to

decide on an initial concept for the aircraft. Three concepts are currently being examined. Since

the aircraft will be a multipurpose vehicle, cabin layouts were designed for combinations of

passenger and cargo payloads. The airplane’s current estimated acquisition cost is $2.5 million,

which is a reasonable price when compared to the Cessna Grand Caravan cost. The GTOW of

the proposed aircraft is considerably more than the comparison aircraft, but the range and

fuselage are both larger than the Cessna Grand Caravan. Future work includes validation of

current sizing and cost estimates, power plant studies, deciding on a final concept, and start work

on the static and dynamic stability of the airplane.

AAE 451 Team 1 page 1 of 36


Table of Contents

1. System Requirements Review………………………………………….………….pg.1 1.1 Product and Market Review……………………………………….…………..pg.1 1.2 Initial Design Requirements……………………………………………….......pg.4 1.3 Design Missions ………………………………………..………………...........pg.5

2. Trade Studies.... ........………………………………………………………….......pg.7

2.1 Engine Selection………………………………………………………………..pg.7 2.1.1 Piston Engines……………………………………………...……………pg.7 2.1.2 Turboprop Engines………………………………………………………pg.8 2.1.3 Comparisons and Conclusions………………………….…………......pg.8

2.2 Fuel Selection.............................................................................................pg.9 2.2.1 Fischer-Tropsch Jet Fuel………………………………………...……..pg.9 2.2.2 Soy-Methyl-Ester-based Jet Fuel………………………….…………..pg.9 2.2.3 Ethanol-based Fuel………………………………………………..…..pg.10 2.2.4 Conclusions………………………………….…………………………pg.10

2.3 Sizing.......................................................................................................pg.10 2.3.1 Sizing Methodology........................................................................pg.10 2.3.2 Trade-off Studies............................................................................pg.12

3. Performance Constraints................................................................................pg.14

3.1 Constraint Diagram..................................................................................pg.14 3.2 Constraint Methodology...........................................................................pg.15

4. Design Selection.............................................................................................pg.17

4.1 Pugh’s Method.........................................................................................pg.18 4.2 Current Designs.......................................................................................pg.20 4.3 Cabin Layouts..........................................................................................pg.22

5. Current Aircraft Definitions and Requirements...............................................pg.29

5.1 Current Characteristics of Aircraft............................................................pg.29 5.2 Comparison with Existing Aircraft............................................................pg.29

6. Conclusions and Further Studies....................................................................pg.30

7. References......................................................................................................pg.31

8. Appendix.........................................................................................................pg.32



1. Systems Requirements Review In this section, the product and target market will be reviewed. In addition, specific

design missions and initial design requirements will be detailed.

1.1 Product and Market Review The product being designed is described as follows: a rugged, dependable, versatile

aircraft powered by non-petroleum based fuel that would serve as a transport vehicle in

underdeveloped and rural areas of the world. Possible customers could be air charters,

governments (land surveying, mail), medical transport, and supply delivery companies that

operate in these rural, developing areas of the world such as Australia, South America, or even

rural parts of Canada.

There are many markets around the world that offer the potential for a successful

business case. Australia and the European Union, for example, have expressed concern over the

environment and thus would be interested in green transportation. There are many international

markets that are also in need of an alternative fuel-based aircraft to replace their existing fleet or

to meet a growing demand.

Figure 1: Population Density of Australia, taken from Swift1

Australia is one of the countries with expressed interest in replacing aging fleets. In a

2003 paper written by S.J. Swift it is pointed out that as a result of the population distribution of

their country; regional airlines are of significant importance to their economy.1 As Figure 1

depicts, most of the population is concentrated in the coastal regions, with 84% of Australians



living on 1% of the continent; the other 16% scattered in rural regions. Because of these

demographics, air transportation is vital to the general welfare of Australia.

The Canadian airline industry also faces many problems, including rapid consolidation,

inconsistent service and rising prices. It is apparent that Canada needs smaller aircraft to service

smaller markets around the country. Between major cities lies vast rural terrain that is not

serviced regularly by the monopolistic national airline Air Canada. According to Dadgostar and

Poulin2, a significant change in the regulation of the Canadian aircraft industry is needed. The

existing market includes several successful regional airlines, including Pacific Coastal Airlines

and Bearskin Airlines. Both of these airlines use existing aircraft that are comparable to the

capabilities and requirements used in developing the current design (see Section 5).

The need for a small, general aviation replacement aircraft is evident in the Latin

American region as well. A very successful aircraft in this market is the Cessna Grand Caravan.

It is ideal for military transport, air rescue, border patrol, surveillance and supply operations.

Cessna Grand Caravans are used extensively by Chile, Brazil, and Colombian government

organizations. Additionally, these aircraft are utilized by airlines in Brazil, Venezuela, and

Mexico. Often serving as an airline transport from remote locations to major airports, these

aircraft offer spacious cabins, room for cargo, and are versatile. An aircraft similar in function

and capacity to the Cessna Grand Caravan is ideal for this region’s needs and has the potential to

compete in a highly successful market.

1.2 Initial Design Requirements Based on market analysis, collected costumer attributes, and an extensive QFD matrix,

the following design requirements were initially considered:

• Passenger range: 9 - 12 • Max Range: 1200 nm • Max Cruise speed: 200 kn • Max Service Ceiling: 25000 ft • Take Off Distance: <2000 ft • GTOW: 10,000 lbs • Acquisition Cost: ≈ $2.0 million (US2006)

Section 5 details current design requirements and target values, which are considerably different

than those listed above.



1.3 Design Missions In order to begin designing an aircraft, initial design missions must be established. All of the

design missions have taken into account the rural routes targeted as well as the lack of

infrastructure needed to refuel in underdeveloped areas. Figure 2 shows a mission profile that is

a contingency plan for an aircraft in the event of a landing being unavailable at the scheduled

airport. This mission would also be used in the event of unfavorable weather conditions or other

extenuation circumstances at the location of the planned landing. The steps of this mission are as

follows:

• 0-1: Take-off (<2000 ft) • 1-2: Climb to <10,000ft • 2-3: Cruise climb <1200nm • 3-4: Descend • 4: Loiter <45min • 4-5: Approach • 5-6: Attempt to land • 6-7: Climb • 7-8: Divert to a neighboring airport • 8-9: Descend • 9: Loiter <45min • 9-10: Approach • 10- 11: Land

Figure 2: Design Mission for Full Range Flight



In this case, the airplane would attempt to land at the prescribed location. Because of

certain circumstances, the aircraft may be unable to complete this step. From here, the aircraft

will climb to an appropriate altitude and divert to an airport within a certain range. The distance

to this neighboring airport is dependent on the location of the original destination. The range of

the diversion is also dependent on the laws in the country of travel. Even though the targeted

market is the international community, most of the planes used in these markets meet FAR 23

reserve requirements, thus the aircraft designed will take into account FAR 23 requirements as

well. From here, the aircraft must descend and loiter again around the airport to which it has

been diverted to.

The mission profile shown in Figure 3 is important in finding the maximum fuel and

thrust necessary to complete the design mission. This specific mission profile represents a flight

from one airport, to a destination, and back to the original airport. The distance from origin to

destination is approximately 600 nm. This is an important mission because the aircraft will be

using alternative fuels, thus refueling stations may not be available at all airports. Thus, the

aircraft will need to return to the original airport in order to refuel for the next mission. The

steps of this mission are as follows:

• 0-1: Take-off (<2,000 ft) • 1-2: Climb to <10,000ft • 2-3: Cruise climb <600nm • 3-4: Descend • 4: Loiter <45min • 4-5: Approach • 5-6: Land • 6-7: Take-off (<2000 ft) • 7-8: Climb to <10,000ft • 8-9: Cruise climb <600nm • 9-10: Descend • 10: Loiter <45min • 10-11: Approach • 11-12: Land



Figure 3: Design Mission for a Round trip Flight without Refueling

The design mission in Figure 3 was taken into account when creating a preliminary constraint

diagram. This profile is discussed in more detail in Section 3.

2. Trade Studies Several trade studies were considered to design an aircraft to meet the initial design

requirements. This research included engine, fuel, and sizing trade studies.

2.1 Engine Selection Fixed wing aircraft currently have four main choices for power plants: turbojet, turbofan,

turboprop, or piston engine. Initial design requirements based on early market analysis showed a

desired cruise speed of 160 knots and 10,000 ft cruising altitude. This ruled out both turbofan

and turbojets due to their inferior thrust specific fuel consumption. Further investigation of

either of the last given choices was necessary for a final decision.

2.1.1 Piston Engines Piston engines for aircraft are available in power capacity ranging from 75 hp to

approximately 500 hp. These engines are capable of running on either Avgas or ethanol as a

replacement. Another fuel option is diesel/kerosene and Soy-Methyl-Ester or Fischer-Tropsch-

based fuel as a replacement. Generally, diesel piston engines for aircraft have a higher specific

weight than Avgas engines but lower fuel consumptions. The same holds true when comparing

piston engines with turbines; refer to Table 1.



2.1.2 Turboprop Engines Turboprops are available in ranges from approximately 400 shp (shaft horse power) up to

10,000 shp, and even higher power output turbines have been in production. They typically have

a lower specific weight than piston engines, and scaling a turbine from a given size to a larger

size is fairly simple. Scaling down, however, is more difficult and results in losses of specific

fuel consumption and a higher specific weight, making small turbines fairly inefficient. Turbines

have a high reliability rate and offer maintenance benefits when compared to piston engines.

2.1.3 Comparisons and Conclusions The team decided to perform a trade study between both kinds of piston engines (diesel

and Avgas) and a turboprop. A sizing code was developed based on turbine-powered aircraft

such as the Cessna Grand Caravan and the Explorer T500. In order to actually compare the

different engines, several changes to the sizing code had to be made to take into account different

specific fuel consumption numbers and weights. This was achieved by linearly scaling numbers

from commonly used engines. The baseline comparison turbine used was the Pratt & Whitney

Canada PT6A-114A rated at 675 shp. This engine was used in both the Explorer T500 and

Cessna Grand Caravan. Comparing this engine to a diesel piston engine (Thielert Aircraft

Engines Centurion 4.0 rated at 350 shp) and an Avgas piston engine (Lycoming IO-720 rated at

400 shp) gave the following results shown in Table 1:

engine type specific weight (lb/hp) sfc at cruise (lb/(hp*h)) empty weight w0 (lb) fuel weight wf (lb)turbine 0.54 0.55 12546 3331Avgas piston 1.43 0.42 14760 3330Diesel piston 1.73 0.36 15396 3209

Table 1: Engine Comparison

Even though both the Avgas piston engine and the Diesel piston engine had a significant

specific fuel burn advantage, their empty weight values outweighed any benefit of specific fuel

values. This fact, combined with the increased acquisition cost and the much heavier airframe

needed for piston engines, lead to the decision to use a turboprop in the aircraft.

The linear scaling of the engines must be questioned. Assuming a power loading value of

0.095 hp/lb, the power required from the piston engines is approximately 1400 hp. There is

currently no piston engine available at even half the power, making the number of engines

necessary to three. Comparing this with a single or twin turbine engine, as well as maintenance

costs and reliability/safety concerns, the decision was made in favor of the turbine engine.



2.1 Fuel Selection When the team decided in favor of a turbine as the engine for the proposed aircraft,

different choices of alternate fuels were considered. Both Soy Methyl Ester based jet fuel and

Fischer-Tropsch based jet fuel can be used in turbines.

2.2.1 Fischer-Tropsch Jet Fuel Fischer-Tropsch (FT) based fuel is almost chemically identical to petrol-based fuel and

can be produced from several different kinds of biomass. Initially explored and used to produce

fuels from coal, the process has been recently modified to be applicable to renewable sources of

biomass such as wood chips, whole corn plants, or other grains.

As described earlier its chemical composition is very similar to petrol based jet fuel.

Several differences include the lack of sulfur in FT jet fuel that eliminates SOx emissions but also

reduces the fuels lubricating quality; this can be counteracted by additives. It also has a slightly

lower energy density than mineral kerosene and thus will have a minimal fuel consumption

increase. This difference can be neglected due to the assumptions used in this stage of the design.

Further research would be necessary at a later stage. Due to its chemical similarity to mineral

kerosene FT jet fuel is broadly compatible with current fuel storage and handling facilities.

2.2.2 Soy-Methyl-Ester-based Jet Fuel Soy Methyl Ester (SMT) fuels have been used and explored over the past 20 years and

are commonly known as Bio-Diesel. Due to gelling problems with SMT fuels at low

temperatures, it was not considered as fuel use in aircrafts. Recent research conducted at the

University of North Dakota3 indicates significant improvements of the low temperature

characteristics thus making SMT a potential aviation fuel.

Dependable figures for SMT performance characteristics in aviation turbines are

currently not available. Figures from piston engines4 indicate a specific fuel consumption penalty

of about 15% due to the different burning characteristics, as well as a lower energy density. Early

trials with fuel blends of regular Diesel and SMT Diesel at different ratios in aviation turbines

performed at Purdue University5 indicate that this is not the case for turbines. However, these

figures are not completely plausible, as they include a number of uncertainties and only consider

fuel blends of SMT with mineral kerosene.



2.2.3 Ethanol-based Fuel Ethanol was also considered as an alternative fuel choice. It was assumed that pure

ethanol and not ethanol blended with gasoline would be considered as an alternate fuel choice.

There are many pros and cons for using ethanol that were considered when making the

alternative fuel choice. Although gasohol, or ethanol blended with a small percentage of

gasoline, is seeing a rise in commercial usage, the use of pure ethanol is not a mature technology.

After considering the positive and negative aspects of using pure ethanol as an alternative fuel,

the decision was made against this type of fuel source. This decision was primarily made due to

the fact that pure ethanol lacks the energy density and therefore the energy potential of other bio-

fuels.

2.2.4 Conclusions Both FT jet fuel and SMT jet fuel are potential alternative fuel sources that are capable of

powering an aircraft turbine. Due to the relative mature state of technology for both fuels and

their general capability of using current infrastructure for storing and handling of the fuel, both

fuels were initially decided to power the proposed aircraft. Due to the lower performance

characteristics of SMT jet fuel and promising FT figures, FT kerosene has been decided as the

alternative fuel for the aircraft.

2.3 Sizing When any product is designed, trade-offs are made to study effects on functionality and

intended use of the product. These trade studies examined the way various flight parameters and

characteristics affected limiting cases of the design requirements. The limiting requirements in

this stage of the design project were payload, range, and takeoff weight.

2.3.1 Sizing Methodology The first step to sizing the airplane was to create a way to study the effects of the design

requirements and mission on the size of the airplane. For this initial sizing, the technique was

taken from the text by Dan Raymer6. This technique used a basic weight buildup equation

(eq.1), and forms of the Breguet Range (eq.2) and Endurance equations (eq.3). With these three

equations, it was possible to create a calculation routine to find a basic weight of the aircraft that

was being designed.



The weight buildup was an iterative process that used several estimated values,

approximated both by the design team and from previously researched aircraft in the database

located in the appendix.

00

0

1WW

WW

WWW

ef

payloadcrew

−−

+=

(1)

Equation 1 was used to create the weight buildup that is discussed in this paragraph. The

first of these was the weight of crew and payload. It was decided that there would be a 2 person

crew on every flight of the design mission. Also, it was determined that each crew member

would be given an average weight of 175 lbs. For each passenger, an average value of 200 lbs

was chosen, which was the standard value used by the FAR 23 requirements. The next portion

of the takeoff weight was the empty weight of the aircraft. This was estimated by taking a

database of similar aircraft and using simple design parameters to create a multivariate

regression for this database. This empty weight equation was found by taking the values of

range, cruise speed, payload, and takeoff weight from the database and using a least squares

method to find an equation that would calculate the empty weight. Finally, the weight of the fuel

was necessary to complete the design mission.

The fuel weight was determined using the Breguet equations for range and endurance.

These equations found the weight change for each leg of the design mission. The design mission

was simplified by having takeoff, climb, and landing sections use average weight fractions taken

from Raymer7 (Table 3.2, pg. 20). The range equation (eq.2) was used to find the fuel

consumption for the cruising section of the design mission.

⎟⎠⎞

⎜⎝⎛

−

−

= DLV

RC

i

i eww

1

(2)

This equation uses the range, C (specific fuel consumption), velocity, and the lift to drag

ratio to find the weight loss fraction for the cruise leg of the flight. The endurance equation

(eq.3) was used for the loiter sections of the design mission.

⎟⎠⎞

⎜⎝⎛

−

−

= DL

EC

i

i eww

1 (3)



The only difference between the range and endurance equation was that the range

equation used distance and speed, whereas the endurance equation used only the flight time. The

L/D of 13.8 was taken from diagrams in Raymer’s6 book, and other values were estimated using

the aircraft database. The specific fuel consumption values were discussed previously in section

2.2 along with the fuel and engine types for the design.

2.3.2 Trade-off Studies Once a way to quantitatively analyze the design choices was developed, the next step was

to compare how changes to design requirements affected other characteristics of the aircraft.

The first set of design requirements given in Section 1 were initially tested with these trade

studies. A secondary observation in all of these studies was the one way trip versus a round trip

that added up to the same total distance.

The first trade study was based on the range of the flight and the payload that was carried

and how that affected the gross takeoff weight. Figure 4 shows this study in the form of a

contour plot of the GTOW as a function of range and payload.

Figure 4: GTOW Trade Study Polar

(Dotted Lines Stand for one way Range, Solid for Round Trip Distance)



For the first trade study, several values were held constant, such as the empty weight of

the airplane. The plane was first sized to complete the round trip design mission. The empty

weight acquired from this initial sizing was then used to determine the takeoff weight of the one

way trip. The empty weight was held constant to determine if the same plane could make both

trips. This was done because the limiting case for this design was the round trip flight of 1200

nm, or 600 nm in each direction. This study found that the original design requirement of a

1200nm round trip with a full load could be completed with the GTOW being below the original

design requirement of 10,000lbs.

The next study determined the most efficient cruise speed. With a constant GTOW, the

cruise speed and range were varied, and available payload was plotted in a contour plot. This

study is depicted in Figure 5. As in Figure 4, the dotted lines stand for a one way trip range and

the solid lines stand for round trip distance.

Figure 5: Payload Study with fixed GTOW, varying Range and Cruise Speed

This contour plot provided the information on cruise speed that was necessary to design

the aircraft. To keep the GTOW less than 10,000 lbs and still meet the original design

requirements, the cruise speed should be kept below 160 knots.



These trade studies provided insight into the possibilities of the intended design. From

the information acquired through these studies, changes were made to the design requirements,

which are discussed in Section 5.

3. Performance Constraints The next step in the design process was to create a constraint diagram, which was

developed by taking various different constraints into account. These constraints were chosen by

the team and will define the performance of the final product. This section details the methods

and equations used to develop the constraint diagrams. Several performance tradeoffs are

defined here.

3.1 Constraint Diagram

0

0.05

0.1

0.15

0.2

0.25

0 10 20 30 40 50 60 70 80

Wing Loading (lb/ft^2)

Pow

er to

Wei

ght R

atio

(hp/

lb)

Sealevel Takeoff

5000 ft Altitude Takeoff on a Hot Day

Sealevel Cruse

10000 ft Altitude Cruse

Sealevel Climb

10000 ft Altitude Climb

15000 ft Cruse

Emergency Landing at 5000 ftAltitude on an Icy Runway

Emergency Landing at 5000 ftAltitude on an Icy Runway With FuelDump OptionEmergency Landing at 5000 ftAltitude on a Hot Day

Comparison Aircraft Design Points

Design Point

Figure 6: Performance Constraint Diagram



3.2 Constraint Methodology The goal of the constraint diagram was to determine the feasibility of an aircraft with

given performance requirements. Through this process, constraints were changed in order to

achieve a feasible design. The diagram compares sea level engine shaft power to weight ratio,

, with wing loading, . Constraints are shown as lines on a graph (see Figure

6). The solution space was defined as the union of the areas of the graph to the left of the

vertical landing constraints and above all the other constraints. A feasible aircraft must have a

and that corresponds to a point in this area. The goal of constraint analysis

was to find the minimum that allows the required performance. Therefore a

smaller, lighter and more economical engine would be possible. The point that

corresponded to the smallest engine along with the GTOW will give the wing area. The final

design point was chosen to be a of 0.095 hp/lb and a of 42 lb/ft

TOSL WSHP / SWTO /

TOSL WSHP / SWTO /

TOSL WSHP /

SWTO /

TOSL WSHP / SWTO / 2. Design

points of comparison aircraft are also depicted in Figure 6. The black point in Figure 6 shows

the vicinity of the chosen design point.

The landing constraints were determined from the time needed for the plane needed to

decelerate to rest after touchdown. The main decelerating force after landing is the braking force

on the landing gear. Aerodynamic forces were neglected in the landing analysis. The braking

force could not exceed the static friction force between the tires and the runway. Otherwise the

plane would skid and lose control. Therefore the runway distance was a function of landing

speed ( )StallV⋅3.1 7 (pg 177) and static friction. The equations in this section were taken from

chapter 5 of Brandt7. Wing loading was solved for to be:

β

μρ⋅

⋅⋅⋅⋅=

69.1max gCs

SW LLTO (4)

Where is runway rolling distance, Ls ρ is air density, is max lift coefficient, g is

acceleration due to gravity,

maxLC

μ is the static friction coefficient, and β is the weight fraction. Since

this equation was independent of engine power, it appears as a vertical line in the constraint

diagram. Feasible designs were to the left of this line because a larger wing area allowed for a

shorter stall speed.



The constraint diagram in Figure 6 shows three landing constraints that allow for a 2000

ft runway. A of 1.8 was used for all the landing constraints, based on comparison aircraft

found in the aircraft database located in the appendix. Density was based on atmospheric tables

for an altitude of 5000 ft. A temperature of 32°F was used for the icy runway landing and 110°F

was used for the hot day scenario. The first two constraints were worst-case scenarios that used

maxLC

μ of 0.1. These constraints differ only in β (fuel ratio), with a value of 0.65, from the sizing

analysis, used to simulate an emergency fuel dump. A μ of 0.6 was used for the dry runway

landing. The two extreme examples of an icy runway were ignored in the final design point

decision since they would require the aircraft to have a massive engine. None of the comparison

aircraft have such strict landing constraints and it was assumed that such adverse weather would

shut down smaller unimproved runways anyway.

At takeoff, the engine produces maximum thrust to achieve takeoff speed ( StallV⋅2.1 )7(pg

173). Takeoff speed is proportional to the inverse of wing area, thus proportional to

Therefore the constraint equation is a straight line with a positive slope:

SWTO / .

S

WgCsW

T TO

LTOTO

SL

⋅⋅⋅⋅=

max

244.1ραβ (5)

Where α represents the thrust lapse and represents the takeoff distance. Propeller aircraft

engine thrust was given in terms of :

TOs

SLSHP

∞

=V

SHPT P

SLSL

ηρρ (6)

Where subscript SL represents sea level conditions and Pη is the propeller efficiency factor.

Solving equations 5 and 6:

S

WgCs

VW

SHP TO

LTOP

SL

TO

SL

⋅⋅⋅⋅= ∞

max

244.1ραβ

ηρρ

(7)

There are two takeoff constraints shown in Figure 6. The only difference between the

two was the air density change caused by the difference in altitude. Alpha α was set at 1 for all

the constraints since the thrust lapse was calculated in equation 6. A of 68 kn was used since

two of the comparison aircraft takeoff at that speed and this value is acceptable by FAR part 23

∞V



standards. A Pη of 0.8 was chosen for all the constraints, as anything higher would be

unrealistic. The takeoff constraints did not affect the final design point.

Other performance constraints come from the equation of specific excess power. Excess

power is equal to the sum of the power devoted to acceleration and the power devoted to climb.

From this:

⎪⎭

⎪⎬⎫

⎪⎩

⎪⎨⎧

++⎥⎥⎦

⎤

⎢⎢⎣

⎡⎟⎟⎠

⎞⎜⎜⎝

⎛+=

dtdV

gdtdh

VSW

qn

SWCq

WT TO

TO

D

TO

SL 11/

2

10 βκ

βαβ (8)

Solving equations 6 and 8 for TOSL WSHP / :

⎪⎭

⎪⎬⎫

⎪⎩

⎪⎨⎧

++⎥⎥⎦

⎤

⎢⎢⎣

⎡⎟⎟⎠

⎞⎜⎜⎝

⎛+=

∞

∞

dtdV

gdtdh

VSW

qn

SWCqV

WSHP TO

TO

D

P

SL

TO

SL 11/

2

10 βκ

βαβ

ηρρ (9)

Where q is the dynamic pressure, the drag coefficient at zero lift, 0DC 1κ the span efficiency

factor, n is the g loading, dh/dt the climb rate, and dV/dt the acceleration. There were four

constraining flight conditions in the plot. For all of these constraints, 1κ of 0.85 was assumed,

based on data from the comparison aircraft. The g loading of 1.3 was used for the constraints

instead of 1 to allow for a safety margin incase of turbulence. The first constraint shown is sea

level cruise of 160 kt. This is placed as a reference line. The 160 kt, 10,000 ft cruise constraint

was the most important since it was based on the design requirements. This line could not be

moved. The 15,000 ft cruise constraint was based on the design requirements. The absolute

ceiling was 15,000 ft since there is no excess power left to provide for climb or horizontal

acceleration at that altitude. The climb constraints allow the plane to climb at 20 ft/s at sea level

and 10 ft/s at 10,000 ft. The original design requirement was for a plane that can climb at 20 ft/s

for its entire ascent. However this requires a larger engine. A climb rate of 10 ft/s is a

compromise to allow for a smaller engine.

4. Design Selection When the aircraft was configured for a specific customer, the placement and selection of

aircraft features and the ability to generate suitable concepts were the most important steps in the

overall design process. Aircraft configurations and designs included such details as engine

placement, passenger comfort, developmental cost, and other details that describe an aircraft.



When design alternatives were selected, it was beneficial for the team to generate as many ideas

as possible. Some ideas were experimental or they were based on limited knowledge, but it was

apparent whether or not a design would be feasible. In the initial design phase, it was extremely

important that the concepts were generated individually. This allowed for individual creativity

and for more ideas to be generated. Once these ideas were put onto paper, the concepts were

selected and individually review by the group. One way of reviewing these generated concepts

was by a selection process known as Pugh’s Method.

4.1 Pugh’s Method Pugh's method is a refined form of a decision matrix method. Selection among listed

alternatives was accomplished by a comparison to a set of criteria defined in the design problem.

Each alternative was weighed by its ability to meet each listed criterion. The criteria selected

aided in choosing and improving the design concepts. The criteria also reflected the potential

customer’s needs and design requirements. This method was used to support judgments about

qualitative information. The method resulted in a satisfactory concept selection for each

alternative configuration.

After the criterion and configuration selections were made an easily readable matrix was

created. A datum concept aircraft was selected to be used as a reference for comparison. Group

meetings were held in order to clarify individual concepts and iterate through the Pugh’s Method

matrix using a system of pluses for positive aspects of different designs, minuses for negative

aspects and “same” (s) for similar criteria. After the ratings were compiled and evaluated,

negatives were improved and positives were enhanced in order to improve the overall design.

The best concept, or the concept that was the most popular, was selected as the new datum and

the matrix was run through again with any new concepts or hybrids. The iterative process of

Pugh’s Method allowed for several concepts to be reviewed and improved based on the selection

of the new datum and running through the comparison matrix.

In the first running of Pugh’s Method, five concepts were compared to the datum, a

Cessna Caravan aircraft, which is shown in Figure 7. The Caravan was selected due to its similar

target customer attributes. Much time was spent as a group going through each individual

criterion for each configuration concept. The results are shown in Table 2.



Criteria

Concept 1

Concept 2

Concept 3

Concept 4

Concept 5

Concept 6

Cessna G

rand Caravan

Maintenance Access - s - - -Cargo Access + s s - sPassenger Space s s s - sPassenger Comfort + s + + +Weight - s - - -Ruggedness - s - - -Conrtolability + s + + sPilot's View area + s - - +Safety s s s - sManufacturability - s - - -Development cost - s s - sWing Design/Configuration + + + + +Tail Configuration + + s s sFuselage Drag - s s - sEngine Placement + s s + +Landing Gea

r + s s - +Plus 8 2 3 4 5Same 2 14 8 1 7

inus 6 0 5 11 4

Datum

M

Figure 7: Three View of Cessna Grand Caravan Table 2: Pugh’s Method Matrix, First Run

After the first run through the matrix, a new datum was selected that was similar to the

Cessna Grand Caravan, but with a few improved characteristics that can be seen in the included

initial Pugh’s Method matrix. This new datum aircraft, which is shown in Figure 8, was then

implemented in the second iteration of Pugh’s Method. Shown in Table 3 are the results.

Figure 8: Datum after First Run Table 3: Pugh’s Method Matrix, Second Run



4.2 Current Designs

run through the matrix, group meetings were held to select the best After the second

concepts. Three vastly different concepts were selected for further research and analysis as seen

in Figures 9-11. The three selected concepts are all unique. Concept 1 is the most similar to the

original target aircraft, the Cessna Caravan. This aircraft is a single-engine, high-wing, turbo-

prop driven plane, which features a large rear cargo door, reinforced landing gear, and a sleek v-

tail configuration. Concept 2 is a twin-engine, high wing, pusher-prop aircraft, which is unique

because is features a quieter and safer cabin due to the engine placement. This design also

includes durable rigid landing gear and a high t-tail. Concept 3, the final configuration being

considered, is a less conventional design which features a tandem wing, which is interesting

because the smaller twin wings work together to reduce induced drag. This aircraft also includes

a high-wing, large rear cargo door, a high t-tail, and durable retractable landing gear.

Figure 9: Concept #1, chosen after two iterations of Pugh’s Method



The three aircraft configurations described will be considered in more detail through

research and comprehensive analysis. There are many remaining factors to consider including

manufacturing concerns, wing design and analysis, and other in-depth trade studies.

4.3 Cabin Layouts The layout of the cabin interior is essential for sizing an aircraft. This aircraft will be used not

only as a passenger transport, but also as a cargo transport. For the cargo aspect of this aircraft’s

mission, the LD-3 industry standard shipping container was chosen. This container has

dimensions of 5’1” by 5’4” by 5’1½”. Because of the designated missions of the aircraft, the

dimensions of the cabin were dependent on three sizing factors. First, the shipping container size

set the cabin height for all possible interior configurations. The width of the cabin was set by

either the width of the shipping container or the width of the seats and aisle. The width of the

cabin was also set by the container in the case when only two seats were placed in a row. In the

case of three seats per row, the width was set by the seat and aisle width. The third and final

constraint on the cabin size was the volume per passenger aboard the aircraft. This constraint

mainly affected the overall length of the cabin. A graph summarizing passenger comfort is

shown in Figure 12.

Figure 12: Passenger Comfort

This graph shows the relationship between passenger comfort, cubic feet per person, and

trip length. The data in this graph was essential in sizing the length of the cabin for certain

flights. For a flight of six hours, a goal of 55 ft3 was set. For shorter flights, 30 ft3 was



acceptable. The pitch of the seats, or the distance between the seats, was also an important factor

in designing the interior of the aircraft. However, this value was simply adjusted to meet the

volume per person goals. In several of the layouts that are to be described, a standard lavatory

has been included in the floor plan, which is 2’10” by 3’2”.

4.3.1 Cabin Size 1 The first layout included a lavatory and five rows of two seats each. Figure 13 depicts

the basic floor plan, including the passenger-only flight option. The overall dimensions of the

interior were 5’6” by 5’2½” by 22’10”.

Figure 13: Two Seat Rows, With Five Rows and a Lavatory

This configuration, as with most of the configurations, could also include the lavatory in

the back of the aircraft; however, this option has not been depicted here. The reason for the

lavatory to be moved to the rear of the cabin is for the comfort of the passengers. Passengers

would prefer not having a lavatory directly across from the entrance. Because this configuration

included a lavatory, it will be used for longer distance flights such as a 1,200 nmi trip. The

cross-sectional area for this configuration is shown in Figure 13. This interior included a

comfortable 26” wide aisle and a 36” pitch. The total interior cabin length was 22’10”. This



gave a comfortable 56.8 ft3 per person. For shorter distances, the lavatory may be removed to

include a sixth row of seats. This caused the pitch to remain at 36”, but each passenger had only

47.3 ft3.

For this size cabin, there were several other configurations that included standard

shipping containers. First, one LD-3 was placed onboard. For this case, there were two main

options for seating aboard the aircraft, with or without a lavatory. With a lavatory, the interior

included eight seats with a 34” pitch and 41.2 ft3 per person, or it included six seats with a 36”

pitch and 54.9 ft3 per person. The other option for one shipping container was without a

lavatory. In this case, there were ten seats with a 35”pitch and 42.0 ft3 per person or eight seats

with a 40” pitch and 52.5 ft3 per person. The second option is depicted in Figure 14.

Figure 14: One LD-3 Container and Eight Seats

An important feature of this configuration is the fact that the cargo is stored in the back of

the cabin, and the passengers are in the front. A cargo door was included near the back of the

fuselage, and a passenger entry door was included in the front of the cabin for this very reason.

This allowed for ease in loading passengers and cargo to their desired locations. Although many

more configurations including shipping containers and passenger seating will be mentioned in

this section, the configurations are very similar to that of Figures 14 and 15, with minor

adjustments to seats per row, and numbers of shipping containers. Thus, not all options will be

depicted here. Another configuration included two LD-3 shipping containers. For this layout,

the lavatory may or may not be included. In both cases the pitch was 36”. However with the

lavatory there was enough space for fours seats with 68.0 ft3 per person. Without the lavatory

there was space for six seats with 45.4 ft3 per person.



The final configuration for this size of cabin included three LD-3 shipping containers.

For this case, there were again two options. The first option is depicted in Figure 15.

Figure 15: 3 LD-3 Containers and a Lavatory

Figure 15 shows a lavatory included for the pilots, in case of longer flights. In this case

there was no room for extra passenger seating. If the flight is shorter, the lavatory may be

removed. For this case, there was room for one row of seats with a 40” pitch. From this, the

passengers had 62.1 ft3 per person.

4.3.2 Cabin Size 2 The second cabin size was very similar to the first lay out; however, the second option

did not include a lavatory in the basic design. This interior had the overall dimensions of 5’6” by

5’2½” by 19’8”. The basic configuration for only passengers can be seen in Figure 16.

Figure 16: Two Seat Rows, with Five Rows Without a Lavatory

This is the basic sizing for the interior of this aircraft. This layout would mainly be used

for shorter distance flights since it did not include a lavatory. For this configuration the

passengers had a 36” pitch, and had a comfortable 47.7 ft3 per person. A longer distance version

of this aircraft can be seen in Figure 17.



Figure 17: Two Seat Rows, With Four Rows with Lavatory

This configuration would be used as a more comfortable flight option for longer trips

since this version of the aircraft included a 40” pitch and 59.7 ft3 per person. One final alteration

to this design was made so that an executive version could be created. This aircraft included a

lavatory again, but only had six seats instead of eight. For this case the pitch was 50” and the

volume per person was 79.6 ft3.

There were three options for this cabin size including shipping containers. These options

were similar to those in the first cabin size; however, all of the options for this cabin size did not

include the lavatory. First, one LD-3 was placed onboard. The interior included eight seats with

a 34” pitch and 41.2 ft3 per person, or it included six seats with a 40” pitch and 54.9 ft3 per

person. Another configuration included two LD-3 shipping containers. In this case the pitch

was 36”. There was enough space for fours seats with 45.4 ft3 per person. The final

configuration for this size of cabin included three LD-3 shipping containers. In this case there

was no room for extra passenger seating.

4.3.3 Cabin Size 3 The third cabin size was a layout that had three seat rows. The interior for this layout had

the overall dimensions of 5’6” by 5’10” by 16’10”. The basic configuration for only passengers

can be seen in Figure 18.



Figure 18: High Density Three Seat Rows, With Three Rows with Lavatory

This is the basic sizing for the interior of this aircraft. For this configuration, the

passengers had a 36” pitch, and had a slightly confined 49.3 ft3 per person. This configuration

was also altered in such a way to allow for more luggage room in the rear of the cabin. This

caused the pitch to decrease to 32”. Of course, this configuration was a bit more confined than

the original layout. For shorter distances, the lavatory was removed and another row of seats

was added. This allowed the pitch to remain at 36”, and the volume per person decreased to 37.0

ft3. Again, this configuration was also altered for extra luggage space so that the pitch was

decreased to 32”.

There were also two other cargo layouts for this cabin size. First, there was the inclusion

of one LD-3 container. For this case, there were two options. First, a lavatory was included.

For this option, six seats were included in the floor plan with a 33” pitch and 46.3 ft3 per person.

This would tend to be a bit confined for longer flights. Without a lavatory, nine seats were

included with a 33” pitch and 30.9 ft3 per person. Since this option did not include a lavatory,

this will be for shorter flights, and this will be a comfortable layout. The other cargo layout

included two LD-3 containers. For this case, there were two options again. A lavatory was

included for longer flights for the pilots and no passengers, or three seats with a 42” pitch and

37.4 ft3 per person without a lavatory was used.



4.3.4 Cabin Size 4 The forth cabin size is similar to the third; however, this is a lower density version of the

three seat row layout. The interior for this layout had the overall dimensions of 5’6” by 5’10” by

18’5”. The basic configuration for only passengers can be seen in Figure 19.

Figure 19: Low density three seat rows, with three rows with lavatory

This is the basic sizing for the interior of this aircraft. For this configuration, the

passengers had a 40” pitch, and had a roomy 55.0 ft3 per person. This configuration was also

altered in such a way to allow for more luggage room in the rear of the cabin. This caused the

pitch to decrease to 36”. This configuration was a bit more confined than the original layout.

For shorter distances, the lavatory was removed and another row of seats was added. This

allowed the pitch to remain at 40” and the volume per person decreased to 41.2 ft3. Again, this

configuration was also altered for extra luggage space so that the pitch was decreased to 36”.

There were also two other cargo layouts for this cabin size. First, there was the inclusion

of one LD-3 container. For this case, there were two options. First, a lavatory was included.

For this option, six seats were included in the floor plan with a 40” pitch and 54.8 ft3 per person.

Without a lavatory, nine seats were included with a 40” pitch and 36.5 ft3 per person. Since this

option did not include a lavatory, this will be for shorter flights, and would be a comfortable

layout. The other cargo layout included two LD-3 containers. For this case, there were two

options again. A lavatory was included for longer flights for the pilots and no passengers, or

three seats with a 60” pitch and 54.4 ft3 per person without a lavatory was used.



5. Current Aircraft Definitions and Requirements After two iterations of Pugh’s Method, trade studies, and developing a satisfactory

performance constraint diagram, the initial design definitions and requirements have changed.

5.1 Current Characteristics of Aircraft Some of the design requirement changes were made according to the information that the

studies provided. Others were altered due to information gained on FAR such as the takeoff

weight. In order to maintain a part 23 certification under the FAR the aircraft must remain under

12,500 lbs Gross Takeoff Weight. This was decided to be a new design threshold for the project.

Design requirements have been updated to the following:

• Passengers: 12 (including crew)

• Cruise Speed: 160 kts

• Takeoff distance: 2000 ft

• Range: 1200 nm

o One-way full load

• Weight: < 12,500 lbs , goal of 12,000lbs

• Cost: < $2.5 million

• Alternate Fuel Source: Fischer-Tropes Kerosene and/or Bio – Kerosene

Current design requirements were decided upon to not only complete the mission that

was set out in the previous requirements review, but also to keep the aircraft competitive in the

market for which it is designed. To do this, it was necessary to comply by FAR part 23

regulations and stay within the price range of aircraft such as the Cessna Caravan and the

Explorer T500.

5.2 Comparison with Existing Aircraft The current aircraft design definitions have several improvements over the Cessna Grand

Caravan. The maximum range of the current design is 1200 nautical miles with a payload of

2,000 lbs, whereas the Grand Caravan’s maximum range is 850 nautical miles at the same

payload. The cabin size for the current design is 22.8’ x 5.5’ x 5.2’ compared to the Grand

Caravan’s cabin size of 16.7’ x 4.3’ x 5.2’; both are designed around standard size shipping

containers so that the plane can be used to transport people as well as cargo. The gross takeoff

weight of the current design is considerably heavier than that of the Grand Caravan, 12,100 lbs

compared to 8,750 lbs. This is partially because the cabin size of the current design is larger than



that of the Cessna Grand Caravan. The wing area and wing span are very similar in size between

the two planes; wing areas are 228 ft2 verses 279.4 ft2. The wing spans are 53.6 ft verses 52.1 ft

for the current design and Cessna Grand Caravan, respectively. The engine power required for

the current aircraft design is 1275 shp, whereas, the Cessna’s engine power required is 675 shp.

Both planes are able to takeoff at many different runways and airports since both planes can

takeoff on runways less than 2,500 ft. Both planes also cost approximately the same amount;

around $2.5 million. These numbers for the Cessna Grand Caravan were taken from the Cessna

Website8. In conclusion, the current design would be a good upgrade and replacement for the

Cessna Grand Caravan and similar planes in the future.

6. Conclusions and Further Studies To conclude, the current design for a non-petroleum fueled aircraft will be a beneficial

addition to the airplane market. This 12 passenger plane propelled by a turboprop engine using

Fischer-Tropsch bio-kerosene will be more environment friendly than the current sources. The

current design will maximize range while considering payload, velocity, takeoff distance, and

cost all at the same time. This product will replace aging, inefficient regional and general

aviation aircraft of its size. Further investigation in aircraft design, dimensions, and cost analysis

will be the next step that will be taken. Specifically, detailed power analysis, static and dynamic

stability, and a more thorough sizing and cost analysis will be explored in greater detail to aid in

the final concept selection.



7. References [1] Swift, S J. Big Challenges For Little Airliners. Australian International Aerospace Congress. Brisbane, Australia: AIAC, 2003. 1-11. [2] Bahram Dadgostar and Bryan Poulin. “Smaller Carriers in Small Markets Better for Customers.” Canadian Business Economics. February, 2001. [3] The Grand Forks Herand: “UND scientists nearing test for new biojet fuel”, http://www.grandforks.com/mld/grandforks/news/13948062.htm [cited 15 February 2006] [4] U.S. Bureau of Mines, Twin Cities Research Center: “Emissions Characteristics of Soy Methyl Ester Fuels in an Underground Mining Diesel Engine with and without Diesel Oxidation Catalyst Aftertreatment”. [5] Melanie A. Kimble-Thom, David L. Stanley, John T. Cholis, Denver W. Lopp: “THE USE OF BIO-FUELS AS ADDITIVES AND EXTENDERS FOR AVIATION TURBINE FUELS”. [6] Raymer, Daniel P. Aircraft Design : A Conceptual Approach, Third Edition [7] Steven A. Brandt, Randall J. Stiles, John J. Bertin, Ray Whitford. Introduction to Aeronautics: A Design Perspective.1997. American Institute of Aeronautics and Astronautics, Inc., Reston, Virginia. [8] Cessna Grand Caravan Performance Specifications:

http://grandcaravan.cessna.com/dimensions.chtml. [cited 9 February 2006]



8. Appendix 8.1 Database



8.2 QFD Matrix



8.3

Sample Concepts for Pugh’s Method



8.4 Tabulated Cabin Layout Specifications

Length Width Height Seats Per Row Cabin Size 1 22'10" 5'2.5" 5'6" 2 Option Lavatory Containers Passengers Pitch Volume Per Person

1 yes 0 10 36" 56.8 ft^3 2 no 0 12 36" 47.3 ft^3 3 yes 1 8 34" 41.2 ft^3 4 yes 1 6 36" 54.9 ft^3 5 no 1 10 35" 42.0 ft^3 6 no 1 8 40" 52.5 ft^3 7 yes 2 4 36" 68.0 ft^3 8 no 2 6 36" 45.4 ft^3 9 yes 3 0 -- --

10 no 3 2 40" 62.1 ft^3 Length Width Height Seats Per Row Cabin Size 2 19'8" 5'2.5" 5'6" 2 Option Lavatory Containers Passengers Pitch Volume Per Person

1 no 0 10 36" 47.7 ft^3 2 yes 0 8 40" 59.7 ft^3 3 yes 0 6 50" 79.6 ft^3 4 no 1 8 34" 41.2 ft^3 5 no 1 6 40" 54.9 ft^3 6 no 2 4 36" 45.4 ft^3 7 no 3 0 -- --



8.4 Tabulated Cabin Layout Specifications (cont.)

Length Width Height Seats Per Row Cabin Size 3 16'10" 5'10" 5'6" 3 Option Lavatory Containers Passengers Pitch Volume Per Person

1 yes 0 9 36" 49.3 ft^3 2 yes 0 9 32" 49.3 ft^3 3 no 0 12 36" 37.0 ft^3 4 no 0 12 32" 37.0 ft^3 5 yes 1 6 33" 46.3 ft^3 6 no 1 9 33" 30.9 ft^3 7 yes 2 0 -- -- 8 no 2 3 42" 37.4 ft^3

Length Width Height Seats Per Row Cabin Size 4 18'5" 5'10" 5'6" 3 Option Lavatory Containers Passengers Pitch Volume Per Person

1 yes 0 9 40" 55.0 ft^3 2 yes 0 9 36" 55.0 ft^3 3 no 0 12 40" 41.2 ft^3 4 no 0 12 36" 41.2 ft^3 5 yes 1 6 40" 54.8 ft^3 6 no 1 9 40" 36.5 ft^3 7 yes 2 0 -- -- 8 no 2 3 60" 54.4 ft^3


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