AAE 451 – Aircraft Senior Design

Spring 2008

Systems Definition Review The Next generation DC-3 Aircraft

Team 4: Skyborne Technologies

Brian Acker

Lance Henricks

Matthew Kayser

Kevin Lobo

Robert Paladino

Ruan Trouw

Dennis Wilde

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Table of Contents

I. Outline 2

II. Mission Statement 3

III. Three Use Cases 4

IV. Major Design Requirements 9

V. Concept Selection 13

VI. Advanced Technologies 18

VII. Cabin/Fuselage Layout 27

VIII. Constraint Analysis 30

IX. Sizing Studies 33

X. Summary 37

Appendix 40

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I. Executive Summary

In the year 2058, the National Aeronautics and Space Administration (NASA),

expects air traffic to have increased significantly. Major hubs all over the world are

expected to be facing congestion and over usage in the future. The problem with

extremely busy hubs is that flights get delayed and passengers miss connecting flights.

NASA asked for the design of a smaller commercial aircraft with the ability to take off

and land in extremely short distances. This new aircraft design is an answer for the

expected problem stated by NASA. The aircraft is expected to carry between 125 and 250

passengers. The takeoff and landing distances need to be less than 3,000 ft.

Skyborne Technologies decided to design a smaller aircraft with a capacity of up to

125 passengers. Based on existing airport data, we found that 80% of the highest traveled

airports are located within 1000 nmi from each other. Using this data, the design range

for this aircraft is 1000 nmi. To solve the problems facing congested airports, two major

factors were incorporated in the design of this commercial aircraft. The first factor is to

utilize smaller airports around large hubs to reduce traffic. The second factor is to

increase the traffic capability of large hubs, by using a split runway approach. The split

runway approach is possible using extremely short takeoff and landing (ESTOL) aircraft

in conjunction with a non-interfering spiral decent. The team at Skyborne Technologies

made use of quick sizing codes based on historic data and found that the takeoff weight

for aircraft of our size is approximately 118,000 lbs. To get this aircraft off the ground in

less than 3,000 ft, several future technologies are being considered. The aircraft will also

have to land in less than 3,000 ft.

The lighter materials and low stall speed allow the aircraft to stop using thrust

reversers, spoilers, and a combination of wheel and speed brakes.

We opted to select a concept so that the sizing would be more specific. A design

process known as Pugh's Method was used to generate concepts and finally to select the

best design. We found that a tandem wing mounted to an elliptical fuselage helps solve

our design requirements. The selection of the concept allowed for some further sizing and

mock up images. Our sizing study expanded with the selection of an airfoil. The airfoil

data made the calculation of key values such as the lift-to-drag ratio and the thrust needed

to take off within the given parameters possible. The fuselage selection allowed for the

design of the cabin layout. The preliminary layout turned out to be uncomfortable

(according to critics), however this is being adjusted. The cabin layout consists of a two

class configuration with a 3+3 seating design for coach and 2+2 seating configuration for

first class. The design will need further analysis to incorporate more future technologies,

which in turn should make the aircraft lighter and more cost effective. Some of the next

steps include the adjustment of values such as the specific fuel consumption and also the

maximum lift coefficient to better represent the expected values for the year 2058.

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II. Mission Statement

“Using innovative solutions to design a short-to-medium range

aircraft that is efficient, eco-friendly, cost effective and capable of

ESTOL, along with new technology to increase passenger traffic.”

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III. Aircraft design missions

Some example design missions are given below.

Mission 1: One way trip up to 1000 nmi

Segments: E-G Optional

A: Taxi/Takeoff (<3,000 ft.)

B: Climb to 35,000 ft.

C: Cruise 1,000 nmi

D: Approach/Land (loiter up to 45 min. if necessary)

E: Climb

F: Cruise 150 nmi to alt. airport

G: Approach/Land

Example:

Travel from Dallas to Orlando, Florida. Replace Dallas Fort Worth to

Orlando International (one of America‟s most heavily traveled routes) with Love

Field to Kissimmee Gateway. Could skip Dallas altogether and choose airport

closer to departure point as well.

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Mission 2: Round trip without refueling (2x 500 nmi legs)

Segments: J, K Optional

A: Taxi/Takeoff (<3,000 ft.)

B: Climb to 35,000 ft.

C: Cruise 500 nmi

D: Approach (loiter up to 30 min.)

E: Land/Takeoff (both <3,000 ft.)

F: Climb to 35,000 ft.

G: Cruise 500 nmi

H: Approach (loiter up to 30 min.)

I: Land (climb to cruise altitude again if necessary)

J: Climb, Cruise 150nmi to alternate airport

K: Land

Example:

1. Round trip from Sacramento, CA to San Diego, CA. Utilize split

runway (need <3,000 ft. for takeoff, landing) to maximize

operations.

2. Cargo carriers such as UPS could use this aircraft to bypass hubs,

flying shipments closer to destination.

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Mission 3: Regional short hops of approximately 300 nmi

Segments: L-N Optional

A: Taxi/Takeoff (<3,000 ft.)

B: Climb to 20,000 ft.

C: Cruise 300 nmi

D: Approach (loiter up to 30 min.)

E: Land/ Takeoff (both <3,000 ft.), Climb to 20,000 ft.

F: Cruise 300 nmi

G: Approach (loiter up to 30 min.)

H: Land/ Takeoff (both <3,000 ft.), Climb to 20,000 ft.

I: Cruise 300 nmi

J: Approach (loiter up to 30 min.)

K: Land

L: Climb to cruise altitude, cruise 150 nmi to alt. airport

M: Approach

N: Land

Example:

1. Travel from New York to Toronto (most traveled U.S. – Intl. route) to

Ottawa to New York. Possible service across Great Lakes Region as well.

2. Rising Populations in Asia could have need for short flights as well

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Justification of Third Mission

Our third design mission involves regional short hops. This expands on the idea of the

second design mission. Our plane will be able to make a multiple flight loop without

refueling. This is possible because our aircraft is designed for a range of 1000 nmi,

although more fuel will be required for three takeoffs the range of the three legs will only

be 900 nmi altogether. We could take-off conventionally within a distance of 3,000ft,

where we can climb and cruise somewhere around 20,000ft and then land within a

distance of 3,000ft. This could be repeated up to two more times with each leg of the trip

somewhere around 300 nmi, including loiter time, climb and approach. This cuts down

the turnaround time and allows servicing more people in more places by not having to

refuel. This will lead to a higher percentage of the plane‟s capacity being utilized. One

possible scenario is from New York City to Toronto, then Toronto to Ottawa and then

finally Ottawa back to New York City. This route of New York City to Toronto is the

highest traveled United States-International flight. This hop might not fully utilize our

capability because we could refuel at each place, but it shows that the airplane can again

make the high volume and demand routes. Experience in traveling near Chicago and

driving around the Great Lake, has shown the process rarely goes smoothly, with

unpredictable weather and extreme traffic congestion. Add in the fact that a straight-line

driving path is impossible and the only available path usually requires considerable

additional driving. A series of quick short affordable flights would be invaluable. Taking

this concept a step further and factoring in the direct point to point service and the ability

to avoid large congested hubs, shows this concept to be viable. Many of the runways in

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this area do not have a full service airport so the ability to make a multi-flight loop

without refueling would be fully utilized.

The multistage hop is ideal for major European cities, due to Europe‟s high

population density and the close proximity of cities to one another. The flights will be

able to connect many more of the European countries, current second-tier cities which

should experience much growth over the next fifty years. China is going through a major

economic boom and will soon require efficient high-speed transportation. Many people

know China is a huge country, but China has an extremely high population density. Most

of the major cities are all located near each other within the range of our aircraft and its

short hop capability. The DC-3 was a must have plane, as it brought the aviation industry

to new places. In keeping with this philosophy, the aircraft design which Skyborne

Technologies is hoping to offer will provide services to locations which currently do not

have them readily available. Even though the plane might not be able to cross the Andes

or the Himalayas, the ability to ferry passengers and cargo over smaller natural obstacles

may be beneficial. South Africa is a great example of a country which has plateaus and

mountains, which make direct routes between cities a hassle. The driving destination

might not be far away but the trip can be extremely long and arduous because of the

unpredictable ground terrain and lack of direct roads. By providing airlines with an

aircraft which can traverse short distances in a cost-effective manner over a short period

of time, Skyborne Technologies will fill a gap in the market.

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IV. Aircraft Design Requirements

a. Meeting Customer Needs

Given the fact that airport congestion and traffic are at all-time highs (and will

increase in the future) smaller airports will become larger players in the passenger

transport service in the near future. Airports may also soon begin to split runways to

perform multiple takeoff/landing cycles on one strip simultaneously. This will allow an

aircraft to takeoff from a runway at the same time that another aircraft is landing which

will allow for more flights to and from some larger airports. Another important fact that

will be discussed shortly is that most people traveling by air fly less than 1000 nautical

miles. Thus, it is likely that airlines will soon be looking for a medium-range aircraft

capable of taking off and landing in very short distances. The new aircraft design will fill

this role perfectly, as it will be capable of taking off and landing in less than 3,000ft and

traveling up to 1,000 nautical miles. This is a non-traditional approach as older airlines

have sought after high range aircraft to allow for flexibility. Our target is a short range

aircraft which agrees with the historical data that will soon be discussed.

Cargo companies may soon be looking for an aircraft with the capabilities

mentioned above to capitalize on split runway configurations and smaller runways. This

aircraft will therefore be offered in a cargo version as well. Although the cargo capacity

of this aircraft (up to 35,000 lbs) less than that of the majority of aircraft operated by

cargo carriers today, short takeoff/landing capability will soon outweigh this capacity due

to airport limitations.

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b. Representative city-pairs

The currently targeted range of our proposed aircraft is one thousand nautical

miles. Thus, transoceanic flights will not be possible with this aircraft. However, most

heavily-traveled U.S. air routes will be within the range of this aircraft. For example, this

aircraft will be capable of flying six of the top ten domestic city pairs for passengers

listed in the figure below (Las Vegas – L.A.; D.C. – Atlanta; Atlanta – N.Y.; Atlanta –

Orlando; D.C. – Chicago; N.Y. – Chicago). The ESTOL capability of this aircraft will

further increase its capacity of these routes (via alternate local airports or split runway

configurations). A system-wide adoption of a short takeoff/landing passenger aircraft

may also lead to less travel on these routes by leading to a move away from the hub-and-

spoke system currently in use today.

Figure 1.1: This figure above indicates distances between the ten most flown domestic city pairs.

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To justify our ideal range based upon the demands for a short range aircraft,

several city pairs were analyzed. To do this we gathered, from the Bureau of

Transportation Statistics, statistics about the top ten most flown city pairs in the United

States shown in Table 1.1.

Table 1.1: Data from the Bureau of Transportation Statistics.

The statistics include the nautical miles between city pairs and the number of

passengers on these flights between November of 2006 and October of 2007. Figure 1.2

below shows a graphical representation of the nautical mileage between these top city

pairs. Notice that all of these cities contain major United States airports and that many of

these ranges are less than a thousand nautical miles.

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Figure 1.2: Millions of passengers over the span of one year.

Based upon this, the percent of passengers flying these top domestic city pairs for

ranges less than eight hundred and one thousand nautical miles was calculated. These

results are displayed up in Table 1.2 below.

Table 1.2: Percent total passengers traveling between the top ten domestic city

pairs based upon nautical miles between November 2006 and October 2007.

We found that 71% of these passengers journeyed less than 800 nmi and 81% of all

passengers traveling between the top ten domestic city pairs journey less than 1000 nmi.

Therefore, our aircraft will be designed for flights of up to one thousand nautical miles.

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V. Concept Selection

The next step for the team was to choose an aircraft concept. We implemented a

concept selection practice known as Pugh‟s Method for this step. For Pugh‟s Method, the

team created several potential aircraft concepts (shown in Figure 1.3). Concepts ranged

from traditional tube-and-wing approach to a flying wing aircraft. All concepts for the

first iteration had fixed wings with the exception of 3, which featured variable-sweep

wings. Concepts 1 and 2 were biplanes featuring two passenger decks that differed in

wing placement. The third design (3) and concept 5 are aerodynamically similar to the

B-2 bomber (3 has a tail and 5 does not) and were affectionately given the designation

„Stingray.‟ Design 4 was a flying airfoil design that was intended to act as a lifting body.

Concept 7 is a tilt-engine approach similar to a tilt-rotor Osprey. Design 8 features

forward-swept wings and fore-mounted Canards while Concept 9 is a twin-fuselage

biplane.

In Pugh‟s Method, each aircraft was given its own column in a matrix. Along the

rows of the matrix, the team listed criteria required of the vehicle to be designed (such as

ESTOL, reliability, and secondary airport compatibility). Team members chose one

concept for use as a „datum,‟ to which each other concept was compared. Concepts were

assigned an „S‟ for a given criteria if they were deemed to have similar characteristics for

that criteria as the datum. A „+‟ or „-„ was assigned if a concept showed more or less

potential respectively for a criteria than did the datum. The first iteration of Pugh‟s

Method is shown below. Design 3, a flying wing with variable-sweep wings, was used as

a datum for the first iteration.

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Figure 1.3: Concepts for first iteration of Pugh’s Method

Figure 1.4: First Iteration of Pugh’s Method

15

As Pugh‟s Method does not use numerical values, the designs were qualitatively

compared using the symbols assigned in the matrix. The first iteration of the method

yielded concepts 5, 4, 6, and 1 as the best. The common tube-with-wings (concept 6) was

discarded, as the team felt it was too conventional and may not be the norm when this

aircraft becomes operational.

Team Skyborne chose the best concepts from the first iteration of Pugh‟s Method

and combined details from the less successful concepts to create new variations. These

concepts were used for the second iteration of the method. For the second iteration,

Concept 1 was a flying airfoil biplane whose wings collapse into one large wing during

flight. Designs 2 and 3 are two-deck biplanes with variable-sweep wings. The only

difference among them is engine placement. Concept 4 is a tandem-wing aircraft that

could feature swing-wing technology if necessary. Design 5 is a biplane with forward-

swept, fixed wings. Design 6 has one wing swept forward and the other swept aft while

concept 7 is a flying wing with variable-sweep wings and no tail. The result of the

second run of the method is shown below.

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Figure 1.5: Concepts for Second Iteration of Pugh’s Method

Figure 1.6: Second Iteration of Pugh’s Method

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The second iteration of the method yielded the two best concepts: a tandem-wing

aircraft with an elliptical fuselage (1) and a flying wing fuselage (4). Due to difficulty

modeling the flying wing as well as its inherent stability issues, the tandem-wing aircraft

was chosen. This design does not include a horizontal stabilizer as the aft wing will be

used for this purpose in addition to its lifting duties. The possibility of swing wings on

this aircraft is available if necessary although Skyborne would like to avoid the need for

this technology due to weight and complexity considerations.

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VI. Advanced Technologies

a. Automated Control Systems

By 2038 unmanned vehicles will have taken over most of the front lines. Unmanned

vehicles are already in service today performing many high risk tasks. The military

already has various unmanned aircraft; the next step would be an unmanned fighter. The

technology is already in use in the military, so why not for commercial transport? Since

the F/A-18 was first designed to be statically unstable, computers have been used to help

pilots to control the aircraft. Computer systems are used in flying wing designs to control

the complex system of controls. Soon it will be possible for an automated system to take-

off and land commercial jets. Such a system can already be implemented but fear over

computer error or outside interference are holding the technology back.

The remote pilot system would be used in conjunction with two other systems to

provide multiple redundancies. These other systems would be having one pilot and one

technician/engineer. The pilot would be fully capable to operate the aircraft while the

technician/engineer would be able to fly the airplane using its automated controls and

computer flying system. The remote system would have to be activated by a person on

board, and then a pilot at the airport or nearby hub could control the plane. This would

prevent the system from being able to be hacked in flight and then a high security

encryption would be used to communicate with the plane during take-off or landing.

This would eliminate a large portion of the plane‟s operating cost. This would cut

airlines‟ pilots cost by about 35 to 40 percent. It would also benefit the pilots because it

would reduce the effects of jet lag and reduce time spent away from their families.

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b. Materials

Titanium has become the premium metal of our age. It offers the all the benefits of

steel while being lighter and less dense. Titanium has only one drawback, its cost.

Cambridge University has recently discovered a process to electrochemically produce

titanium. The FFC Cambridge process reduces titanium oxides to produce a pure metal

alloy with much less energy and is much simpler than the currently used Kroll process.

The Kroll process uses lots of energy to melt the metal and is very complex. It takes a

few days for just one batch of titanium. Another new development for titanium is

powdered titanium. The United States government is funding the development.

Titanium powder would allow the creation of complex shaped parts; this would reduce

the need for machining the parts. This would drastically reduce the amount of wasted

metal and the cost of machining the part to its desired shape. Metal prices double about

every 15 years, but with all the new developments for titanium its price will increase

slower and double closer to 30 years. This will allow commercial aircraft to stop using

heavy steel components. Aluminum still has one advantage of titanium, its weight and

will continue to be used in the fuselage skin of future airplanes.

Composite materials are quickly becoming the material of the future. They are

the ideal solution. Composites are stronger and lighter than there metal comparisons.

Carbon fiber is 6 times stronger than aluminum with a weight savings of 40 percent.

They also can offer better resistance to both heat and vibrations. They also require less

maintenance. Composites will probably never fully replace all the metal on an airplane,

but they can be used in conjunction with metals to further strengthen them. Composites

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have already been used in the Boeing 787 and are planned for the new Airbus 350. They

can even be used in critical structures like the wings. Both companies are claiming a 14

to 15 percent decrease in empty weight. Carbon nanotubes are the newest and most

promising composite material. Carbon nanotubes are almost 70 times as strong as

titanium and half the density of Aluminum. They have already been used to reinforce the

winning Tour De France bicycle by using extremely small carbon fibers woven together

to reinforce other materials and metals. New manufacturing processes are also helping to

cut the empty weight of the aircraft. One piece fuselage sections are reducing the number

of rivets and fasteners. In the future these sections will be able to be built bigger and

longer reducing the number of rivets and fasteners even more. The following graphs

illustrate composition by weight.

Skybourne Technologies

75%

7%

15%0% 3%

Composites

Aluminum

Titanium

Steel

Other

Skyborne Technologies

Fig 1.7: Material Composition of Skyborne’s Concept (By Weight)

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This will result in approximately a twenty-one percent total weight savings when

applied to our concept. The twenty-one percent calculation comes from multiplying the

two weight savings factors of .86 and .92. These factors come from the original weight

savings of a fifty percent composite single aisle and the second from the reduction of

steel and addition of more composites along with new manufacturing processes. Our

aircraft has two wings which can both use composites, but a smaller fuselage because

these are estimates for wide body aircraft.

B-787

50%

20%

15%

10%5%

Composites

Aluminum

Titanium

Steel

Other

A350

52%

20%

14%

7%7%

Composites

Aluminum

Titanium

Steel

Other

Fig 1.8: Material Composition of Boeing 787 (By Weight)

Fig 1.9: Material Composition of Airbus A350 (By Weight)

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c. Variable Sweep

Variable sweep was one of the original solutions to problem of the increased drag

at higher speeds. The graph below shows this problem by illustrating how the ideal wing

sweep changes as the Mach number is increased. The second graph shows how the huge

increase in drag can be offset by sweeping the wing proving this systems benefit.

Fig 1.10: Optimal Wing Sweep and Wing Drag Coefficient v. Mach

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Variable Sweep has been used on many successful aircraft such as the Mirage G,

B-1, F-111, and F-14. It does have a few drawbacks. It has heavy complex parts that

require more maintenance and add to the empty weight of the aircraft. Historically, it

adds about 4 percent to the empty weight. New technology advancements have allowed

for it to be considered for use on a commercial aircraft. The old swing wing systems

were designed with slide rules, not computers. The F-14 was designed extremely quickly

and was a huge improvement on the original variable geometry wings. The titanium box

that houses the pivot point could be reinforced with carbon nanotube decreasing weight

while increasing strength. The large hydraulic worm drive that pivots the wing could be

replaced by a system of smaller actuators. This system of actuators could provide the

redundancy needed for FAA certification and reduce weight. Even though it could work

for a commercial transport it may not be needed with our tandem wing design. We have

enough lift at slower speeds and high sweep angle because of our large wing area. Also,

we do not take advantage of its full benefit because our required sweep is only around 25

degrees and not the standard 60 degrees. Further analysis will determine whether or not

variable sweep wings will be employed on the aircraft. At the present time, variable

sweep wings are estimated to add 5.5 percent of the aircraft‟s empty weight to the overall

design weight.

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d. Variable Incidence

Variable incidence in the past has been used to provide an increase in the angle of

attack (AoA) at take-off and landing. This increase in AoA leads to an increase in the lift

coefficient. It is achieved with similar components as variable sweep, which add weight

and complexity. We first explored it as a way to achieve more lift, but now it is only a

consideration to help alleviate stability problems caused by a tandem wing design with

possible variable sweep wings. It has been used successfully in the Martin XB-51 and

the F-8. Further analysis will be performed to determine if variable incidence wings will

be put into use on the new aircraft design.

e. Propulsion

Unducted fans were predicted to be 35 percent more fuel efficient while producing

less pollution than conventional turbojet engines. Unducted fans saw a 30 percent

decrease in specific fuel consumption in the 1980‟s. With the current state of this

technology it was found to be just a bit impractical. Unducted fans are capable of Mach

0.75, but produce high cabin noise. At the present time, our engines will be mounted to

the fuselage and not the wings, so cabin noise is a factor of great concern. When landing

at smaller airports, noise will be a huge concern for the surrounding community. Also if

a blade severs from the engine the damage done to the fuselage could be catastrophic.

These disadvantages might be overcome in time, but at too high of cost. There are other

turbofan or turbojet power sources available which may provide the same specific fuel

consumption and efficiency as the unducted fan.

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Since unducted fans might not be practical a conventional turbofan that burns

biofuels might be. With the rising oil costs and environmental concerns biofuels would

help to alleviate both of the two major problems.

Current turbofan advancement is happening at an alarming rate. The turbofans on

the original 747 produced around 25,000 lbs of thrust. When the engines were updated

20 years later the new engines produced 5,000 more pounds of thrust and saved 14

percent more fuel. The GeNx program took only 10 years to replicate the same amount

of progress, a savings of 15 percent off of the specific fuel consumption. The decrease of

specific fuel consumption over the last fifty years is charted in Fig. 1.11 below. Based on

the graph, a specific fuel consumption value of 0.3 has been used for design analysis.

Fig. 1.11: Turbofan SFC v. Year

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Reusable energy sources like solar power could be harnessed to provide power to

run the compressor or power the fuselage. Getting rid of or drastically reducing the size

of the APU would cut the costs and weight of the aircraft. Powering the compressor with

electrical energy not coming directly from the engine would greatly increase the specific

fuel consumption. The energy extracted by the turbine stage of the engine could be

reduced and more of the power from combustion can be extracted. The solar panels and

batteries required to hold the charge would add weight and complexity, but it could

greatly reduce operating costs and environmental impact. Solar power is already in use

and has proven to be effective.

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VII. Cabin/ Fuselage Layout

Figure 1.12: Current Cabin Layout

The final layout design is shown in the picture above. One item of note is that except

for the fuselage, the position and size of the aircraft components are subject to change.

The location of the center of gravity and aerodynamic center are vital to determining the

stability of the aircraft. More research is being done to determine the optimum location

for the wings and engines. As shown, the concept consists of a standard tube body but

with tandem wings to provide enough lift for ESTOL capabilities. There is a vertical tail

but a horizontal tail is not necessary as the aft tandem wing will fulfill that function. At

this point both the fore and aft wings are capable of variable sweep. At takeoff the wings

are unswept to provide the maximum coefficient of lift. In cruise the wings are swept

back in order to decrease the drag and increase the cruise flight speed. However,

variable sweep wings are historically very heavy and expensive. While an argument can

be made that such costs will be decreased, a determination is being made as to whether

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having variable sweep wings is necessary. The variable sweep would provide a shorter

takeoff distance but a compromise may be reached where a fixed sweep wing can

maintain the threshold value while reducing the drag in flight. As an added benefit the

engines could then potentially be mounted on the fore wing thereby getting rid of the

extra support necessary for fuselage mounting.

An important aspect of the design concept is the fuselage. The shape and size of the

fuselage has dramatic impacts on the weight and aerodynamic performance. It was

decided to design an aircraft with only a single aisle. Judging by current aircraft and the

number of passengers it was deemed unnecessary to have more than one aisle. The

seating is divided into first-class and economy sections with total passenger seating for

126, not including the crew. For the most recent configuration there are three rows of

first class seating with four seats per row. The economy section is comprised of nineteen

rows with six seats per row which is more than the usual 2+3 seating of many

commercial single-aisle aircraft. The result of this is to produce a shorter aircraft which

can translate to weight savings. The original cabin shape chosen was a constant area

cylinder. This was done for simplicity and pressurization purposes. A circular cross-

section fuselage is very simple to manufacture. A circle is also easiest to pressurize as

there are no stress concentration points from changing geometry. However, the price is

paid in passenger comfort. A circular cross-section means that the shape of the fuselage

will intersect with the space of the person sitting at a window. From the comments given

it was acknowledged that for the given timeframe of the concept it is safe to assume that

there will be no structural issues from pressurization. Also, it can be argued that

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advanced manufacturing will eliminate the advantage that a circular fuselage currently

has. To provide more comfort to the passengers the cabin shape was changed to an oval.

The benefit of the new shape is to provide a more comfortable flight for window seat

passengers along with increased aisle height and more head room. The downside is that

the maximum diameter of the aircraft has increased, increasing the weight and parasite

drag. It still needs to be determined what diameter will provide adequate passenger

comfort without increasing the size too much. The current specifications of the aircraft

layout are:

Cabin length: 70 ft

Maximum diameter: 13 ft

Fuselage Length: 115 ft

Table 1.3: Layout Specifications

Seat width (in) Seat Pitch (in) Aisle Width (in)

First-Class 24 39 24

Economy 17 34 18

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VIII. Constraint Analysis

In order to determine the acceptable design range for the aircraft, a constraint diagram

(Fig. 1.13) was created. The figure was created by inputting a range of wing loading

values (Wo/S). A range of thrust to weight (T/W) values were then determined. The

results were plotted for various sectors of the aircraft flight profile. The areas of greatest

concern were: takeoff, climb, initial cruise, cruise, descent and landing. The following

formula was used to determine the thrust-to-weight curve for takeoff:

𝑇𝑆𝐿

𝑊0=

1.1 2𝛽2 𝑊𝑜

𝑆

𝛼𝑔𝜌 𝐶𝐿𝑀𝐴𝑋𝑆𝑇𝑂

(Equation 1.1)

The desired value for the takeoff distance, STO, was less than 3000 ft. The CLmax

value was input from the result generated via the sizing code.

To approximate curves for climb, initial cruise, cruise and descent, variations on

the following formula were used:

𝑇𝑆𝐿

𝑊0 =

𝛽

𝛼 𝑞

𝛽

𝐶𝐷𝑜

𝑊𝑜

𝑆

+ 1

𝜋𝐴𝑅𝑒 𝑛𝛽

𝑞

2

𝑊𝑜

𝑆 +

1

𝑉

𝑑ℎ

𝑑𝑡+

1

𝑔

𝑑𝑉

𝑑𝑡 (Equation 1.2)

In Equation 1.2, for initial cruise and cruise, values for both dV/dt and dh/dt were

set equal to zero. Based on the desired Mach number, values for dynamic pressure, q,

were determined. Values from the table of aircraft weight fractions were used to

determine various values for β. For descent, the desired rate of descent at this time was

found to be (-25) ft/s.

For landing, Equation 1.3 was used to determine the wing loading. During

landing, thrust-to-weight should be kept to a minimum, as the goal is to slow the aircraft

31

as quickly as possible. The desired landing distance, SL, was selected to be less than

3000 ft. The braking coefficient, μ, was set to be 0.5, based on data from Boeing. Based

on Equation 1.3 below, a constant value for (Wo/S) was generated. This was then plotted

against a range of T/W values, to produce a vertical line.

𝑊𝑜

𝑆=

𝜌𝑔𝜇 𝑆𝐿𝐶𝐿𝑚𝑎𝑥

1.15 2𝛽 (Equation 1.3)

Once curves for takeoff, climb, initial cruise, cruise, descent and landing had been

generated, the results were plotted on the same graph. For takeoff, climb, initial cruise,

cruise and descent, the required thrust-to-weight values were determined to be those

greater than the generated curve. For landing, the required thrust-to-weight values were

those to the left of the vertical line. Based on this, Figure 1.13 was generated. Shading

accordingly, the green sector of the graph represents the acceptable thrust-to-weight

values for a given range of wing loading values. From the constraint diagram, the

maximum wing loading value was determined to be 96.33 lbs/ft2. Based on the sizing

code, the current wing loading for the aircraft is 87.7 lbs/ft2. Thus, the wing loading is

within the acceptable portion of the constraint diagram.

32

Fig. 1.13, Constraint Diagram

33

IX. Most Recent Sizing Studies

a. Sizing approach

Skyborne Technology‟s sizing is based off an airfoil similar to the one used on the

Boeing “Dash 80” commercial aircraft, found on an airfoil database maintained by David

Lednicer1. The decision was made to use the NACA 64211 airfoil (Fig. 1.15), following

research on wind tunnel data using a java applet created by Dr. Martin Hepperle2. This

airfoil shows a good lift-curve slope (Fig. 1.16), which gives a high coefficient of lift for

an appropriate Reynolds number at 35,000 ft (Fig. 1.17). This 2-D case was then used to

calculate many performance characteristics of a 3-D model. The assumptions made when

determining values for variables were based on historical data, such as an Oswald

efficiency factor of 0.8 and our takeoff gross weight estimate. An aspect ratio of around

11 was used as opposed to the historical data of about 9.5 due to our tandem wing design.

Matlab was used as the preferred sizing approach, along with scripts created using

textbook equations, those provided in AAE 451 lectures, (see script in appendix and a

quick-sizing spread sheet). Further analysis deemed this to be an efficient sizing

approach, taking into consideration the time constraints and futuristic technology which

were in place, such as our tandem variable sweep wings. This sizing study assumed a

rubber engine design, which was hard to analyze via usage of detailed sizing packages

such as FLOPS or ACS. A disadvantage of the sizing approach is the inability to analyze

characteristics such as noise profiles and carbon footprints, which are important to take

into consideration.

34

Using Matlab, the „Big Six‟ design/performance variables were determined. For

the aircraft, which can carry 126 passengers with a gross takeoff weight of 87,700 lbs, a

cruise 𝐶𝐿𝑀𝐴𝑋 of 0.98 was used, with a fuel fraction of 0.146 and a cruise Mach number of

0.80. The lift to drag (L/D) ratio for cruise is 19.6, which is fairly reasonable for a

commercial aircraft of this size. The thrust to weight ratio at cruise is 0.17 and the wing

loading is 87.72 lbs/ft2, without taking into account the use of any high lift devices. The

desired range of around 1000 nmi was achieved while staying within all of the design

constraints. These results are favorable for capabilities such as ESTOL and efficiency.

However, further analysis of the sizing code in order to better suit customer needs will be

required. These needs would include accommodating a higher level of passenger comfort

and decreasing the specific fuel consumption (SFC). These needs will be analyzed after

benchmarking improvements in advanced technology and projecting them into the future.

In order to evaluate the effectiveness of the aircraft design as compared to the

previously stated goals, a compliance matrix has been implemented. Currently, this

matrix shows that we meet most of the stated target values and are thus satisfied with the

design at this point (see table 1.3).

35

Sizing Summary (Cruise Conditions)

Passengers - 126 @ 220 lb/pax

Crew - 5

Operating empty weight (We) – 46,400 lbs

Gross takeoff weight (Wo) – 87,700 lbs

Fuel weight (Wf) – 12.800 lbs

Empty weight fraction – 0.528

SFC – 0.30 1/hr

Fuel Fraction – 0.146

AR – 11

Wing Sweep – 32°

Vcruise – 778.5 ft/s (461 kts)

M @ 35,000 ft – 0.80

CLmax = 0.98 (Cruise)

L/D – 19.63 (Cruise)

T/W – 0.167 (w/o high lift devices)

Wing Loading (W/S) – 87.72 lb/ft^2

Range – 1,000 nmi

36

Sizing Summary (Takeoff Conditions)

Passengers - 126 @ 220 lb/pax

Crew - 5

Operating empty weight (We) – 46,400 lbs

Gross takeoff weight (Wo) – 87,700 lbs

Empty weight fraction – 0.528

SFC – 0.30 1/hr

Fuel Fraction – 0.146

AR – 11

Wing Sweep – 32°

Vstall – 186.0 ft/s (110.2 kts)

Vtakeoff – 223.3 ft/s (132.3 kts)

CLmax – 2.13

T/W – 0.245

Wing Loading (W/S) – 87.7 lb/ft^2

37

X. Summary

Skyborne Technologies is well on its way to designing the next generation short-

to-medium-range airliner. This aircraft will be lightweight, reliable, and cost-efficient.

To help solve airport congestion problems of the future, the aircraft will be designed

primarily with ESTOL capabilities in mind. The aircraft will utilize a range of no less

than 1,000 nmi. Airlines and cargo carriers will be the primary customers for our aircraft.

The aircraft will be designed with several different missions in mind, including direct

flights between major city pairs and shorter regional hops.

Fig. 1.14: Isometric, Top View of Concept

Skyborne Technologies presented multiple concepts for this aircraft. A tandem-

wing aircraft with an elliptical cross-section design was selected, using a process known

as Pugh‟s Method. The current layout is shown in the picture above. The position and

size of the aircraft components are subject to change. Research is being done to

38

determine the optimum location for wings and engines. Current sizing sees this aircraft

carrying 126 passengers and 5 crew at Mcruise of 0.8 with a range of 1,000 nmi. Gross

takeoff weight is currently 87.700 pounds.

Unducted fans saw unprecedented decreases in specific fuel consumption in the past

few decades and are thus being considered for this aircraft. However, noise concerns

may make this technology impractical. Advances in titanium production technology may

make this metal economically practical by the time this aircraft enters service.

Composites are already in use in airliners but this aircraft will take that utilization to a

whole new level, allowing for up to forty percent decrease in empty weight. Carbon

nanotubes are seventy times as strong as steel and may well be an option when this

aircraft enters production.

There are several steps that need to be taken from here. A more accurate sizing will

be done using updated SFC and 𝐶𝐿 coefficient values to minimize necessary weight and

wing area. A decision will be made as to whether variable-sweep wings will be

necessary or not. The center of gravity and aerodynamic center will be estimated in order

to make sure that our aircraft is statically stable. A decision must be made as to whether

or not static stability will even be an FAA requirement when this aircraft enters service.

Following these steps should ensure that Skyborne Technologies remains on track to

accomplish its mission.

39

References

1Lednicer, David and Selig, Michael, UICC Airfoil Database - The Incomplete Guide to Airfoil

Usage, Analytical Methods, Inc, Redmond, WA, 98052, 2007.

2Hepperle, Martin. JavaFoil©, <http://www.mh-aerotools.de/airfoils/javafoil.htm>, Myrtenweg

1, D-38108 Braunschweig, Germany, 1996-2006.

3Bureau of Transportation Statistics. January 31, 2008. < www.bts.gov >

4Jane‟s All the World‟s Aircraft. January 20, 2008. < jawa.janes.com/docs/jawa/search.jsp >

5National Aeronautics & Space Administration. February 2, 2008. www.nasa.gov

6Ott, James. “Combating Congestion.” Aviation Week & Space Technology. Jan 7, 2008: 41.

7Raymer, Daniel P. Aircraft Design: A Conceptual Approach, 4th Edition. Reston, VA.

8American Institute of Aeronautics & Astronautics, Inc., 2006.

9Rolls-Royce Corp. “Market Outlook 2007.” < www.rolls-royce.com > 10Brandt, Steven A. “Introduction to Aeronautics: A Design Perspective.”2nd Edition. Reston, VA..

40

APPENDIX

Figure 1.15 – NACA 64211 Drag Polar

Figure 1.16 – NACA 64211 Lift-Curve Slope

41

Figure 1.17 – Laminar flow-field around the NACA 64211 Airfoil

Table 1.4 – Compliance Matrix

Target Threshold Current

Runway Length (ft) <2000 <3000 <3000

Range @ max payload (nmi) 1000 800 1098

MCRUISE 0.82 0.78 0.8

Ramp weight (lbs) - - 118000

Price (M $) 40 45 -

Passengers 125 150 126

42

Matlab Script (Cruise Conditions)

clear all,close all,clc

%%%%%%%%%%%%%%%%%%%%%%%%%%

%Wing Geometry

%%%%%%%%%%%%%%%%%%%%%%%%%%

S_for = 600; %ft^2

S_aft = 400; %ft^2

S_tot = S_for + S_aft; %ft^2

AR_for=11; % Assumed due to tandem wing design

AR_aft=11;

b_for=sqrt(AR_for*S_for); %ft

b_aft=sqrt(AR_aft*S_aft); %ft

c_for_mean=S_for/b_for ; %ft

c_aft_mean=S_aft/b_aft ; %ft

sweep_for=32; %degrees % Picked to fit the 0.8 Mach #

sweep_aft=32; %degrees

e = 0.8; %acceptable value for oswald efficiency

g = 32.2; %ft/s^2\

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

%Initial design weights

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

xvec = Wo_iterator; % Function that iterates the take of weight

wo = xvec(1) %GOTW, lbf

wf = xvec(2) %Weight of fuel, lbf

we = xvec(3) %OEW, lbf

cdo = 0.015; %Typical Coefficient of Drag at zero lift

SFC = 0.3; %Specific Fuel Consumption

Beta = 0.97; % Mass fraction for take-off

alpha_ltr = 1;

% atmosphere

rho_0 = 0.002377; %Density at 35,000 ft, slug/ft^3

Viscocity_0=0.3745e-6; %Viscocity at 35,000ft slug/ft s

% 2-D case

cl_alpha_for=0.1076; %/degree based on wind tunnel data

cl_alpha_aft=0.1076; %/degree based on wind tunnel data

alpha_Lo_for=-1.8; %alpha L=0 based on wind tunnel data

alpha_Lo_aft=-1.8; %alpha L=0 based on wind tunnel data

alpha_Lmax_for= 9.8; %alpha Lmax based on wind tunnel data in

degrees

43

alpha_Lmax_aft= 9.8; %alpha Lmax based on wind tunnel data in

degrees

% High Lift Devices

chord_extension = 0.1; %Chord extension with chord = 1

zerolift_deflection = -10; % degrees

dCL_max = 0.9*1.9*(1+chord_extension)*cos(zerolift_deflection*pi/180);

% 3-D case

CL_alpha_for=cl_alpha_for/(1+((57.3*cl_alpha_for)/(pi*e*AR_for)));

%/deg pg131 reference 10

CL_alpha_aft=cl_alpha_aft/(1+((57.3*cl_alpha_aft)/(pi*e*AR_aft)));

%/deg pg131 reference 10

CL_for_max=CL_alpha_for*(alpha_Lmax_for-alpha_Lo_for) + dCL_max

%pg 123 of reference 10

CL_aft_max=CL_alpha_aft*(alpha_Lmax_aft-alpha_Lo_aft)

%pg 123 of reference 10

CL_max= ((CL_for_max*S_for)+(CL_aft_max*S_aft))/(S_for+S_aft) %for same q,

didnt take into account downwash, will need to add q's when eda is found

%%%%%%%

% The CL_max was weighted against the planform area of each wing

%%%%%%%

% Stall speed

Vstall = sqrt((2*wo)/(rho_0*CL_max*(S_for+S_aft))) %

Reference 7

%Take off Velocity

Knots_ft_s = 1.6878;

V = 1.2*Vstall %ft/s

Vk = V/Knots_ft_s %knots

Vstall_knots = Vstall/Knots_ft_s %knots

%reynolds number

Re_for=(rho_0*V*c_for_mean)/Viscocity_0;

Re_aft=(rho_0*V*c_aft_mean)/Viscocity_0;

%CL cruise required

q = (1/2)*rho_0*(V^2); %slug/ft s^2

Drag

=(cdo*q*S_tot)+(((CL_for_max^2)*q*S_for)/(pi*e*AR_for))+(((CL_aft_max^2)*q*S_

aft)/(pi*e*AR_aft)) %lbf

Cd_ac =Drag/((q*S_for)+(q*S_aft)) %Aircrafts drag coefficient; note,

when eda is found need to revise with different q's

% Take-off Distance

Distance = 2500; %ft (ground roll)

Lift = CL_max*q*(S_tot)

Thrust = (1.44*wo^2)/(rho_0*S_tot*CL_max*g*Distance) + Drag +

Viscocity_0*(wo-Lift)

44

%Aircraft Specifications

L_D = CL_max/Cd_ac

%L/D for max alpha at TO

W_S = wo/(S_for+S_aft)

%Wing Loading total, lb/ft2

T_W_TO = ((1.1)^2*(Beta)^2*(W_S))/((alpha_ltr)*g*rho_0*CL_max*Distance)

%Thrust to Weight Ratio for max alpha at TO

Range=(Vk/SFC)*(L_D)*log(1.119727539) %Breguet Range Equation for

Cruise

45

Matlab Script (Wo Iteration)

function [xvec] = Wo_iterator()

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

% Empty weight Prediction

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

Woguess(1) = 0;

Woguess(2) = 21000; % Initial guess

AR = 11;

T_to_W = .275;

W_S = 95;

M_max = 0.8;

Beta = .14633;

Wpay = 27500;

Wcrew = 1000;

weight_saving = .9;

k = 2;

while Woguess(k) > 1.0001*Woguess(k-1) || Woguess(k) < .9999*Woguess(k-1)

We(k) = (1.7766*(Woguess(k)^0.1399)*(AR^0.0923)*(T_to_W^0.1829)*(W_S^-

0.5685)*(M_max^0.4260))*Woguess(k)*weight_saving;

Wf(k) = Woguess(k)*Beta;

k = k+1;

Woguess(k) = We(k-1) + Wf(k-1) + Wpay + Wcrew;

end

Wo = Woguess(k);

Wfuel = Wf(k-1);

Wempty = We(k-1);

xvec(1) = Wo;

xvec(2) = Wfuel;

xvec(3) = Wempty;