Table of contents

EXECUTIVE SUMMARY.................................................................................................................................1

MISSION STATEMENT..................................................................................................................................2

AIRCRAFT DESIGN MISSION.........................................................................................................................2

SELECTED BEST AIRCRAFT CONCEPT............................................................................................................3

ADVANCED TECHNOLOGIES................................................................................................................7

RESULTS OF AIRCRAFT SIZING AND CARPET PLOTS.....................................................................................9

MAJOR DESIGN TRADEOFFS......................................................................................................................15

AIRCRAFT DESCRIPTION.............................................................................................................................16

AERODYNAMIC DESIGN ANALYSIS.............................................................................................................25

PERFORMANCE..........................................................................................................................................32

STRUCTURES..............................................................................................................................................35

WEIGHTS AND BALANCE............................................................................................................................45

STABILITY AND WEIGHTS...........................................................................................................................46

NOISE.........................................................................................................................................................51

SUMMARY.................................................................................................................................................63

REFERENCES..............................................................................................................................................66

EXECUTIVE SUMMARYWhile the technology employed in business aviation is ever advancing, the

ultimate goal of comfortably and quickly transporting corporate customers to their

required locations remains steadfast. World-conscious airframe designers must not only

be aware of their customer’s needs, but their environmental impact as well. Team two has

designed a business jet aircraft capable of meeting our target customer’s needs while

maintaining environmental responsibility. The main requirements of a long range, high

cruise speed, low environmental noise and low emission aircraft were established in the

System Requirements Review document. A number of concepts were generated to meet

the requirements, and two designs were chosen in the Systems Definition Review

document. Finally, one design was selected which best met the requirements, which was

optimized for its design mission during this document, the Conceptual Design Review.

In order to meet the customer needs and NASA’s Environmentally Responsible

Aviation Project N+2 goals, it was determined that drastic changes would need to be

made to modern business jet aircraft. In an effort to meet those goals, advanced

technologies were analyzed on a lifting canard aircraft with a swept back wing

incorporating two geared turbofan engines.

A sizing code was developed, and historical and advanced concept data was

researched. Aircraft weights and balance, aerodynamics, engine selection, and noise were

analyzed in the design process. Major structural component locations were investigated,

the noise produced by the aircraft was predicted, and a cost analysis has been conducted

for our aircraft. The aircraft’s performance was also determined, and consideration was

given to the major design tradeoffs used throughout the design process.

Our aircraft is capable of transporting 16 passengers in a luxury cabin with a

maximum still air range of 7,100 nm. Our aircraft will weigh 71,300 lbs, and requires

3,900 feet of ground roll. All of the customer needs were met or exceeded, but the NASA

N+2 goals proved too difficult to meet with our design. Though not obtaining the NASA

N+2 goals, our aircraft provided a 20db reduction in environmental noise, a 50%

emissions reduction, a 25% fuel burn reduction, and a 33% takeoff field length reduction

over modern aircraft of the same class. While unsuccessful in reaching the NASA N+2

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goals, the customer needs were met; significant environmental improvements were

achieved, and a great deal was learned about the aircraft design process.

MISSION STATEMENTThis project’s main goal is the design of a cost effective, high speed, luxury

aircraft capable of transporting customers to any destination as quickly as possible. The

project’s secondary goal is to meet NASA’s N+2 criteria, reducing the environmental

impact of the aircraft. Our proposed aircraft will be able to compete with other aircraft of

the ultra long range category.

AIRCRAFT DESIGN MISSION The design mission was developed and optimized with the city pair of Los

Angeles and Hong Kong in mind, accounting for a 60 knot head wind. Eight mission legs

between nine points constitute the developed design mission as illustrated in Figure 1

below.

Figure 1: Design Mission Flight Plan.

The first leg of the mission, from points 0 to 1, is taxi and takeoff to an altitude of

50 feet. Points 1 to 2 represent the climb portion of the mission where the aircraft climbs

at best rate of climb to an altitude of 41,000 feet. From there the aircraft enters the cruise

leg of the mission, between points 2 and 3, and begins cruising at a Mach number of 0.85

2

for 6350 nautical miles. Cruise is then followed directly by a no-range credit descent to

land where the aircraft will attempt a landing from points 4 to 5. In the absence of a

landing, the aircraft will then climb to an altitude of 5,000 feet at best rate climb, depicted

between points 5 and 6, and cruise at an altitude of 5,000 feet to an alternate airport 200

nautical miles away. Upon arrival at the alternate airport, the aircraft will enter a holding

pattern for 45 minutes, from points 7 to 8, and then begin a no-range credit descent to

land. Finally, the aircraft lands at the alternate airport and completes the last mission leg

at point 9.

While the design mission is the intended use of this aircraft, several other

operating missions can be made by this aircraft as well. One additional operating mission

example would be flying from New York to Van Nuys. The distance between the two

cities, which is 2,146 nautical miles, falls well within the design mission range of 6,350

miles. To compensate for the largely unused range, the aircraft can then be flown at its

maximum Mach number of 0.9 with a maximum capacity of 16 passengers. This range

tradeoff allows for tremendous flexibility in speed and capacity for shorter ranged flights.

SELECTED BEST AIRCRAFT CONCEPTMotivated by our mission statement, the system requirements design stage produced a set

of criteria by which individual concepts could be evaluated based upon the customer needs.

Following the initial system requirements design stage, eight concepts were generated for further

consideration. Pugh’s method was implemented as an objective procedure for selecting the best

aircraft to narrow the eight concepts down to two. Early conceptual sketches of the aircraft are

shown in Figure 2, concept one on the left, and concept two on the right. Some of the design

criteria used in Pugh’s Method were the NASA N+2 goals, which are listed below in Table 1.

Figure 2 Aircraft Concepts Selected by Pugh’s Method

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Table 1 NASA N+2 Goals Technology Benefits Relative

To a Large Twin Aisle Reference Configuration

Corners of

the Trade SpaceN+2 Goals

Noise(cum. below Stage 4)

-42 dB

LTO NOx Emissions (below CAEP 6)

-75%

Performance:Aircraft Fuel Burn

-50%

Performance:Takeoff Field Length

-50%

Concept two possessed a potential reduction in induced drag at cruise due to the

incorporation of a lifting canard, thereby addressing a reduction in aircraft fuel burn.

Additionally, concept two addressed noise reduction by offering a main wing and twin vertical

tails positioned to shield engine noise from the aircraft’s surroundings. NOx emission reduction

was tied directly to aircraft fuel burn, thereby offering a potential edge to concept two due to its

lower cruise drag. Neither of the two designs displayed a clear advantage in performance field

length.

After evaluating each of the two concepts through Pugh’s method, the potential to satisfy

the NASA N+2 goals, and sizing studies, the canard configuration from Figure 2 was selected for

the remainder of our design.

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Aircraft Walk-around

Several important features provided the canard design with advantages over the other

aircraft configurations evaluated in the concept selection process. Particular aspects from all or

many of the concepts were also shared with the canard configuration’s final design in order to

meet the specified criteria. A number of the unique and shared external characteristics present

on the canard configuration aircraft are depicted in Figure 3 below.

Figure 3 Canard Concept Walk-Around Chart

Shown in Figure 3, the canard design features lifting canards in an effort to obtain a

reduction in induced drag during the cruise portion of the design mission. Also aimed at

reducing induced drag at cruise, our design incorporates spiroid wing-tip devices as shown. The

previous features were included to fulfill the aircraft fuel burn reduction metric of the NASA

N+2 goals listed in Table 1. These features also act to reduce NOx emissions since a lower fuel

burn will produce fewer emissions.

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Lifting Canard

s

Fuselage – aft Mounted Engines

Noise ShieldingVertical Stabilizers

Noise ShieldingLow Wing

Circular Fuselage

Spiroid Wing-Tips

Also present on the canard configuration are several noise reducing design features,

which are shown in Figure 3. Engine noise was determined to be the primary contributor to

environmental noise, and the canard configuration made several efforts to shield the environment

from noise generated by the engine. The wing was placed low and aft on the aircraft, beneath

both of the engines to shield the engine noise from the ground during the takeoff, climb, descent,

and landing mission segments. As lateral noise propagation is also a significant concern for

airports in populated areas, twin vertical stabilizers were incorporated on top of the main wings

outboard of the engines, and in line with the engine exhaust.

From the walk-around chart depicted in Figure 3, it is clear that several of the external

design features directly impact the internal cabin of the aircraft. Both engines are mounted aft of

the main cabin and on pylons connected to the fuselage. The aft mounting of the engines affords

a quieter cabin interior, while the fuselage mounted pylons facilitate a reduction in cabin

vibration and provides a minimum structural weight for engine mounting.

The design of a circular fuselage cross section provides a relatively easy method of cabin

pressurization, and allows a relatively large internal cabin volume for passenger comfort. A

circular fuselage can be constructed from current industry practices, thereby lowering

manufacturing developmental costs.

The low, aft mounted main wing also allows for minimal main spar and cabin interaction.

The lack of structural requirements inside of the main cabin helps maximize internal cabin

volume, yielding additional passenger comfort. Furthermore, the aft mounted main wing

facilitates a clear view of the ground from nearly every seat in the cabin, which provides each

passenger with a visually pleasing exterior view.

The benefits of the aircraft’s design features from Figure 3 provide insight to the reasons

for the choice of concept. A more thorough investigation of the important design features and

advanced technologies incorporated on the design follows in the next section of the report.

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ADVANCED TECHNOLOGIESOf all the advanced technologies that could be included, engines are perhaps the most important

since the engines will have a significant impact on all four of the NASA N+2 goals. Our group

considered two engines, an unducted turbofan and a geared turbofan. For reasons discussed later in the

report, a geared turbofan was chosen over the unducted fan. A geared turbofan engine utilizes a

gearbox between the fan and main power shaft to decouple the fan from the engine’s compressor.

The technology allows for a very large fan to rotate at slower speeds while moving large

amounts of air with little noise. The gearbox also allows for the compressor and turbine to spin at

faster speeds to further increase efficiency. Pratt and Whitney claims that the PurePower 1000

series offers proven efficiency with no life limited parts, a 20 dB reduction in noise over current

engines, a reduction in NOx emissions by 50% over the CAEP/6 margins, and a 15% reduction

in fuel burn over current engines.3 The PurePower 1000 series offer a thrust range between

13,000 lbf and 24,000 lbf depending on engine selection. This thrust range is ideal for our

aircraft and allows the team to continue using the rubber engine assumption in sizing.

Another advanced technology included was the use of composite materials.

Aircraft such as the Boeing 787 have achieved a significant empty weight reduction by

utilizing composite materials in the majority of the airframe. A reduction in aircraft

weight will allow for reduced fuel consumption and a reduction in take off length. Up to

50% of the structure of some aircraft has been constructed out of composite materials.

There has been a lot of research on the use of composite materials within the past few

decades. The technology has been proven to be reliable and beneficial in the aerospace

industry and is currently being fielded in many production designs. This technology is

likely to keep improving and delivering even more advantages in aircraft design by the

2020 production date. Research and past applications suggest a reduction of

approximately 20% in structural weight. This value is used directly as a technology

factor in the sizing process and is applied to the structural weight components of the

fuselage, wings, horizontal and vertical tails, canards, pylons, and nacelles.

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Spiroid wingtips were the last advanced technology incorporated into the design.

These wingtips have been shown to significantly reduce fuel burn in flight testing. When

the wingtips first flew on a Gulfstream II, they yielded a 10% improvement in fuel burn.

Figure 4 shows the general geometry of these devices, which estimates state can provide

a 6-10 % reduction in cruise drag. This technology has been incorporated on more than

3,000 aircraft, including several business jets, as well as the Boeing 737 and 757 airliners.

Spiroid wingtips also aid the FAA in increasing airspace capacity near airports. Spiroid

wingtips have the potential to decrease wake intensity, which could substantially alter the

requirements for separation distances between aircraft in airport traffic patterns.

Figure 4: Spiroid Wingtip Devices

Values of Major Design Parameters

With the basic aircraft configuration, important design features, and advanced

technologies chosen, the design was iterated through the developed sizing code until the

aircraft’s major design parameters were optimized for the design mission. A number of the

optimized parameters including thrust to weight ratio, aspect ratio, wing loading, and main wing

and canard areas were selected and are shown in Table 2. While each of these parameters can be

further broken down into dimensions more commonly provided for an aircraft, these dimensions

are more easily shown in an aircraft three-view drawing provided in the Aircraft Description

section of this report.

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Table 2: Major Aircraft Design Parameters

Parameter Value

Thrust / Weight Ratio 0.34

Aspect Ratio 12

Wing Loading 87 (lb/ft2)

Wing Area 796.4(ft2)

Canard Area 147.4 (ft2)

RESULTS OF AIRCRAFT SIZING AND CARPET PLOTSIn order to numerically evaluate and size our designs, a sizing code was

developed in MATLAB. The code follows the general procedures outlined in the flow

chart shown in Figure 5.

Figure 5: Sizing Code Flow Chart

The empty weight calculation is based on a component build-up method using

equations from the Transport Aircraft category of Daniel Raymer’s textbook.8 Once the

code was established; it was calibrated against the Beechcraft Starship for our canard

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design and against the Gulfstream G550 for our conventional design. The results of the

calibration process yielded the calibration factors listed in Table 3.

Table 3: Calibration Factors

Weight Conventional Canard

Fuel Weight 0.89 0.89

Empty Weight 1.16 0.96

Gross Weight 1.03 0.98

It is important to note that the fuel weight calibration factor is the same for both

designs, which was intentional. In calibrating the canard model to the Beechcraft

Starship, the fuel weight calibration factor was extremely small, as low as 0.50. This

large discrepancy was attributed to the turboprop engine used on the Starship. Because

of the fundamental differences between a turboprop and a geared turbofan or unducted

fan, the fuel weight calibration factor found for the conventional design was used to

replace that of the canard design. Finally, with the code completed and calibrated, the

technology factors were then applied and each model was analyzed. A list of the

technology factors used in the sizing code is provided in Table 4. The value of each

factor comes from the applied advanced technologies discussed previously in this report.

Table 4: Technology Factors

Application Tech Value

Wstructure 0.80

Di (canard only) 0.93

SFC 0.75

The conventional and canard aircrafts were sized and optimized through the use of carpet

plots by running the sizing code through many iterations by varying thrust to weight, wing

loading, and aspect ratio. These three variables were the main driver for sizing the aircrafts. The

most optimized case was a result of whichever combination of these variables gave the lowest

gross weight. The goal of this optimization was to produce the lightest plane that could carry out

the design mission. Several constraints limited the weight of the aircraft and the three design

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variables. The constraints for the aircraft were top of climb (>100 ft/m climb rate at cruise

altitude), 2-g maneuver, takeoff ground roll (<4000 ft), landing ground roll (<3000 ft), and climb

gradient for second segment climb (2.5%).

Thrust to weight and wing loading were varied in order to find the minimum gross weight

at a given aspect ratio. Five different aspect ratios were tested for each model, and a trend was

then established between minimum gross weight and aspect ratio. Below in Figure 6 is the trend

that was seen for the conventional aircraft. From this plot it can be seen that the minimum gross

weight for the conventional design is 76,000 lbs at an aspect ratio of 10.

8 8.5 9 9.5 10 10.5 11 11.5 127

7.2

7.4

7.6

7.8

8

8.2

8.4

8.6

8.8

9x 10

4

W0

AR

Aspect Ratio vs W0 for Conventional a/c

Figure 6: Gross Weight vs. Aspect Ratio of Conventional Aircraft

Further analysis on this design point can be seen by examining the carpet plot for

an aspect ratio of 10, shown in Figure 7 below. If a point lies within an area shaded in

grey, it means that it cannot meet one or more of the constraints. Thus the minimum gross

weight is at a wing loading of 85 and a thrust to weight of .32 which yields a gross weight

of 76,000 lbs. This point is limited by both top of climb and takeoff ground roll. Landing

distance, second segment climb, and subsonic 2-g turn are not limiting constraints.

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Figure 7: Carpet plot for conventional aircraft

The same analysis was done for the canard aircraft. A similar trend was seen for

the relationship between gross weight and aspect ratio. A plot of this trend is shown in

Figure 8 below. The major difference from the conventional design is that the optimal

aspect ratio is now slightly higher at 12 compared to 10. At an aspect ratio of 12, the

aircraft weights only 71,000 lbs. This is nearly 5,000 lbs lighter than the conventional

aircraft and is one of the reasons for choosing the canard configuration over the

conventional.

12

10 10.5 11 11.5 12 12.5 13 13.5 146.5

7

7.5

8x 10

4

AR

W0

Aspect Ratio vs W0 for Canard a/c

Figure 8: Conventional a/c sizing results

The carpet plot for the canard aircraft corresponding to the minimum gross weight is

shown below in Figure 9. This shows that the minimum gross weight occurs at a wing loading of

approximately 87 lb/ft2 and thrust to weight slightly below 0.34. This optimal point is

constrained by top of climb and takeoff ground roll. The other three constraints were not a factor,

which was the same result as the conventional aircraft.

13

Figure 9 Carpet plot for conventional aircraft

From this sizing optimization, the results show the canard aircraft is approximately 5,000

lbs lighter than the conventional design. The canard aircraft weights 71,300 lbs, of which 38,000

lbs is empty weight and 31,500 lbs is fuel. The takeoff ground roll is just below 4,000 ft and the

landing ground roll is 2,200 ft.

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MAJOR DESIGN TRADEOFFSOne of the concepts selected proposed the use of a lifting canard as opposed to a

conventional tail configuration. The lifting canard can theoretically provide the design

with better cruise performance by reducing induced drag at cruise conditions. In a

conventional tail configuration the horizontal stabilizer provides a down force in flight

for aircraft stability, therefore the wings have to create additional lift to counter act the

down force from the tail which in turn means the wings create additional induced drag.

The lifting canard balances the pitch stability of the aircraft along with providing lift,

therefore the main wings do not have to produce as much lift as a conventional design

would require. The cons for a canard can outweigh the benefits. Specifically, the

downwash from the canards could actually produce enough interference on the main

wings to actually create more drag. In fact, the interference between the canard and main

wing occurs in almost all flight conditions except for the specific design point at cruise

where the canard configuration will benefit induced drag. Additionally, the FAA

mandates that the canard must stall before the main wing to prevent loss of aircraft

stability. Should the main wings stall before the canards, the lifting force provided by the

canards will cause a violent pitch up causing the main wings to go into a deeper stall and

possibly prevent the aircraft from recovering. The pros of having the canards stall before

the main wing actually prevents the main wings from stalling unintentionally by

preventing the aircraft from flying at high angles of attack. The cons of having the canard

stall first means that main wings cannot ever reach maximum lift which requires the

addition of high lift devices for slow flight and landing. The position of the main wings at

the very rear of the aircraft actually provides a difficult design problem for installing high

lift devices due to the large pitching moment created by high lift devices so far aft of the

center of gravity and careful attention must be paid when designing the aircraft for

landing configurations.

The canard design was calculated to have a smaller empty weight than the conventional

design, but also had a slightly larger fuel burn during the design mission. Therefore the canard

design actually had a worse drag profile than the conventional aircraft which would suggest that

the canard did not perform well in reducing the aircraft’s drag during cruise. The team decided to

use the canard configuration despite this fact to try and meet the N+2 noise reduction goals. The

15

canard position allows the engines to be placed in the rear of the aircraft above the main wings

which provide a significant amount of noise shielding to the ground. Two vertical stabilizers are

also mounted on the main wings of the aircraft to reduce engine noise. The two engines are

essentially blocked by the main wings and vertical stabilizers which significantly reduced the

noise signature of the aircraft.

Additional tradeoffs were considered pertaining to the aircraft’s cabin. Space

versus performance considerations for the cabin involved installing two lavatories, a

galley, and larger sleeping quarters for a reserve pilot. The addition of such amenities

forced the cabin size to be rather large, but a large selling point in business aircraft is

comfort. The two lavatories were necessary for a full passenger load, especially for long

distance flights. The galley was added for long distance flights including trans-pacific

flights where meals would have to be served due to the long duration. The pilot resting

quarters were added for a reserve pilot, which in the case of the trans-pacific flights a

reserve pilot is mandated by FAA law due to limits on pilot-in-command time. The cabin

also was designed to have a stand up center aisle to push the design into the “plush”

category of aircraft comfort.

AIRCRAFT DESCRIPTIONOnce the external design features of the aircraft were determined, the aircraft

configuration was iterated through the sizing code, the major design parameters were optimized,

and the aircraft’s outer mold line was set and modeled in CATIA V5, R19. The model

incorporated the design features as previously mentioned, a low rear wing design with a lifting

canard, circular fuselage, twin vertical outboard mounted tails, and aft-fuselage mounted engines

on pylons. The fuselage seats sixteen passengers, two crew, two pilots, and an additional pilot in

the pilot rest area for trans-oceanic missions at maximum capacity. The following three

subsections offer a more detailed investigation of the aircraft’s outer mold line, the aircraft’s

interior layout, and the included cabin amenities and features respectively.

Dimensioned Three – View To Scale

From the aircraft sizing code and the major design parameters listed in Table 2, a

thorough set of dimensions were calculated for the aircraft concept, which are shown in the scale

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three-view drawing of the aircraft presented in Figure 10, generated from the CATIA model.

The three-view drawing was constructed at a scale of 1:300 with reference to the actual aircraft’s

completed size when plotted on standard 8-1/2” x 11” paper.

17

Figure 10: Dimensioned Scale Three View of Team 2 Final Design, Scale 1:300

18

To provide additional ease of reference, a table of the major exterior aircraft dimensions

was compiled and is presented below in Table 5. The values previously listed in the major

design parameters compilation presented in Table 2 are omitted here for brevity.

Table 5: External Aircraft Dimensions

Internal Layout / Arrangement To Scale

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Dimension Value

Main Wings

Wingspan 97’-9”

Wing Sweep 30˚

Root Chord 13’-9”

Tip Chord 2’-8”

Strake Sweep 45˚

Strake Root Chord 24’

Canards

Wingspan 36’-5”

Canard Sweep 35˚

Mean Chord 3’

Vertical Tails

Span Each 13’-5”

Tail Sweep 30˚

Mean Chord 10’-1”

Width Between 30’-6”

Fuselage

Overall Length 88’

Main Cabin Length 50’

Main Cabin Outer Diameter 8’-10”

Sill Height 6’-3”

With the dimensions of the outer mold line and arrangement in place, the interior

arrangement of the aircraft could be finalized. The crew area of the aircraft including the flight

attendants reclining seats, as well as the pilot berthing area were located at the foremost part of

the main cabin, forward of the main entrance door. Seating for sixteen passengers at maximum

capacity was distributed about the main cabin, with a conference area, lavatory, and small

cocktail galley present in the main cabin area. An emergency exit was also placed across from

the conference area and over the leading edge of the main wing on the opposite side of the

aircraft from the main entrance. At the rear end of the main cabin area, a divider wall separates

the rear lavatory, preparation galley, and in cabin storage from the main seating area. As with

the external mold line of the aircraft, the cabin interior was completely modeled in CATIA V5

R19, and a 1:150 scale three view drawing was generated and is depicted in Figure 11.

Figure 11: Overall Interior Cabin Three-view Drawing

In order to provide more insight to the interior cabin dimensions, a detail drawing

of the conference and small galley area was generated from the above plan, and more

dimensions were added, as shown in Figure 12. Additionally, a cross sectional view of

the interior and exterior cabin dimensions was constructed to provide a clearer view of

the relevant seating dimensions and aisle height, and is depicted in Figure 13.

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Figure 12: Detail Drawing of Conference Area and Cocktail Galley

Figure 13: Fuselage Cross Section and Seating Dimensions

Similar to the table constructed for the exterior aircraft dimensions, a table of interior

layout and seating dimensions was constructed for the main cabin, shown in Table 6.

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Table 6: Interior Cabin and Seating Dimensions

Dimension Value

Major Cabin Dimensions

Outer Cabin Diameter 106”

Inner Cabin Diameter 100˚

Aisle Height 6’-3”

Minimum Aisle Width 27”

Main Cabin Length 50’

Additional Cabin Length 10’

Seating Dimensions

Standard Recliner Width 27”

Standard Recliner Depth 33”

Square Table Length / Width 26”

Flight Attendant Recliner Width 20”

Flight Attendant Recliner Depth 33”

Sofa Dimensions

Sofa Length 7’-6”

Sofa Depth 35”

Conference Area Dimensions

Conference Seat Width 4’-4”

Conference Seat Length 8’-4”

Conference Table Width 2’-6”

Conference Table Length 4’-4”

22

Cabin Layout / Amenities

The aircraft’s cabin was designed with the intent of providing a plush experience for each

and every passenger on board. A number of amenities and features were incorporated to

heighten the travel experience. The main cabin features a modern technology conference area,

complete with four passenger conference seating around a common table, a conference computer

desk and chair, and a retractable projector screen. More typical aircraft seating can be found in

the six large fully reclining chairs, two pairs of which are arranged in a face to face seating and

share common fold down tables. For more casual seating, two sofas are included in the main

cabin area, capable of seating three passengers each with a maximum capacity, or one passenger

each in complete luxury. An emergency exit is provided on the opposite side of the main cabin

entrance at the rear of the cabin and over the leading edge of the main wings, near the conference

area as shown in Figure 14.

Figure 14: Emergency Exit Located at Rear of Main Cabin

As the business traveler requires sustenance to maintain a high level of productivity, two

galleys were included in the aircraft cabin layout. A large fully equipped galley is located at the

rear of the aircraft designed for both flight attendant and passenger use; while a smaller more

23

easily accessible cocktail and serving galley is located near the center of the aircraft, and near the

conference area, shown in Figure 15.

Figure 15: Cocktail and Serving Galley, as Viewed from Conference Area

Based upon the maximum cabin capacity, it was ascertained that two full lavatories

would be required to provide the comfort level desired. One lavatory was placed forward in the

fuselage near the crew seating and main entrance door, and one placed after the rear main cabin

divider easily accessible from the conference area.

Cabin design and comfort was not only focused around the passenger, a comfortable crew

rest area was incorporated in the main cabin, isolated from the passenger area. A curtained pilot

berthing area was incorporated in the foreword most section of the main cabin, capable of

providing a comfortable rest area for a third pilot in rotation on extended flights as shown in

Figure 16.

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Figure 16: Pilot Berthing Area and Crew Rest Seating

Additionally, two fully reclining flight attendant seats are located forward in the main

cabin, also capable of being separated from the main cabin seating area by a dividing curtain.

Storage for the cabin materials and crew baggage was provided by the reserved areas aft of the

rear cabin divider adjacent to the main preparation galley. A rendered overhead view of the

internal cabin is provided below in Figure 17.

Figure 17: Overhead View of Internal Cabin Layout

At full capacity, the aircraft cabin holds a maximum of sixteen passengers, two crew, and

three pilots. Also at maximum capacity, each passenger is provided with a volume per passenger

of 150 ft3. The volume per passenger metric equals or exceeds other plush business aircraft of

the same class, but is typically specified in terms of volume per passenger per hour, from which

varying comfort levels are classified. These comfort levels are defined in Figure 18.3

Extrapolating the data to longer durations and larger cabin volumes, it can be easily shown that

the 150 cubic feet offered by the cabin design maintains the plush level of comfort for any trip at

maximum capacity, and improves even further for shorter missions.

25

Figure 18: Comfort Level of Cabin Volume / Passenger / Trip Duration3

Given the large internal cabin volume, the plush comfort level for any duration and

capacity flight, and ample amenities, it becomes fundamentally clear that our interior cabin was

designed with every level of passenger and crew comfort and productivity in mind.

AERODYNAMIC DESIGN ANALYSISAirfoil Selection

To meet the desired lift on the aircraft, the airfoils for the main wing, canard, and

the vertical tail were selected based on the needed coefficient of lift, or CL, for three

different phases of flight. These phases were the takeoff CL, cruise CL, and landing CL.

These three values are listed in Table 7.

Table 7: Desired Coefficient of lift

CL

Takeoff 1.2

Cruise 0.46

Landing 2

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From the sizing code, the airfoil was set to have a 0.1 thickness to cord. This limited the

airfoil selection some. The supercritical wing is advantageous for this aircraft as it provides lift

and stall characteristics comparable to non-supercritical wings while minimizing wave drag as is

inherent in the design of a supercritical wing. It was also ideal to pick a supercritical airfoil

because it would lower the MDD for the drag prediction. To pick a supercritical airfoil, the RAE

2822 was used as a baseline model. This was done because the RAE airfoil had much more

information available for it. Airfoils with the proper thickness to cord were compared using CL

vs. angle of attack plots. The results of this comparison can be seen in Figure 19. This

comparison was done using a relatively low Reynolds number of 100,000. It is important to note

that the hump that each of the airfoils has is due to the fact that the information is being gathered

at a relatively low Reynolds number. From this table, the NASA SC(2)-0610 airfoil was selected.

This particular airfoil was selected because it can meet the desired cruise CL of 0.46 with having

minimal twist in the wings. This particular airfoil has a significant pitching moment, but it was

deemed to be manageable and could be worked around if the lifting canard could balance this

out.

Figure 19: Supercritical Airfoil comparison13

27

With the airfoil selected, it can be analyzed at different Reynolds numbers. The

geometry information for the NASA SC(2)-0610 is given in Figure 20. A visual of the

airfoil can also be seen in Figure 20, and more information about the geometry can be

seen in Table 8. Using this geometry, the airfoil had been further evaluated at different

low Reynolds numbers. This comparison can be graphically seen in Figure 21. It is

important to note that as the Reynolds number increases, the stall angle transitions higher

along the curve. The drag polar for this airfoil can also be seen in Figure 22, where the

trend is that as the Reynolds number increases, the coefficient of drag decreases. The

canard used a slightly thinner supercritical airfoil, with a thickness to cord of 0.08. This

thinner airfoil, NASA SC(2)-0608, does not provide as much lift or as great of a moment,

but is more than sufficient to meet the requirements of the canard. This airfoil will also

stall earlier depending on the angle that it is set at, which will help to keep the aircraft

stable.

Table 8: NASA SC(2)-0610 geometry

Geometry

Thickness: 10.00

%

Camber: 1.00

%

Trailing edge

angle:

2.9

deg

Lower Flatness: 12.50

%

Leading edge

radius:

1.70

%

28

Figure 20: NASA SC(2)-0610 airfoil.

Figure 21: CL and Cm vs. Angle of Attack13

29

Figure 22: Drag Polar13

The selected airfoil meets the cruise CL, but does not meet the climb or the landing CL of

1.2 and 2.0. To obtain the required CL for these two phases of flight, a high lift device needs to

be included on the aircraft. The landing CL was used for choosing the appropriate high lift device

because it is the higher of the two coefficients of lift, requiring a CL of 2.0. To find the effects of

adding a regular or slotted flap, historical data from Kenneth W. Goodson’s report was used to

see how greatly the two types of flaps would affect a supercritical airfoil.14 The airfoil used for

this data is different than the airfoil selected for our aircraft, but it provides a good baseline for

comparison of the effects the flaps will have on the coefficient of lift. The results from this data

are shown in Figure 23. From this plot it can be seen that the regular flap will not produce a high

enough CL to reach 2.0. The single slotted flap, on the other hand, can reach the desired amount

of lift with only a minimal angle of attack, assuming the flap is deflected at a sufficient angle.

30

Figure 23: Effects of flaps on a supercritical airfoil14

The next airfoil that needed to be selected was that of the vertical tail. For this, a laminar

flow symmetric airfoil was ideal. This was picked because the vertical tail has a relatively small

operating range, and does not need to produce lift at a zero-degree angle of attack. Laminar flow

airfoils extend the transition point from laminar to turbulent flow on the airfoil. This can help to

reduce drag. The desired thickness to chord was obtained from the sizing code and found to be

0.08. After comparing different laminar flow airfoils, the NACA 64(1)-008 was selected. This

airfoil can be seen in Figure 24. The resulting Cm vs. angle of attack for this airfoil can be seen

in Figure 25. This data is again taken at a very low Reynolds number. When a rudder is attached

to the airfoil, it will generate the desired Cm needed for maneuverability.

31

Figure 24: NACA 64(1)-008 Laminar flow airfoil.

Figure 25: Cl and Cm vs. Angle of Attack for Vertical Tails13

Drag Prediction

Throughout our design we considered drag to be built up of four components,

namely wave, induced, pressure and miscellaneous drag. We also considered the

miscellaneous drag to be a steady five percent of the pressure drag. This meant that we

immediately reduced the number of drag variables from four to three. Our earlier reports

simply assumed the wave drag coefficient to be 20 “counts,” or CD,wave = 0.002. Our

pressure drag approximations were much more in depth and took into consideration all of

the major components of the aircraft including the fuselage, engines, nacelles, pylons,

32

wings etc. The remaining, or miscellaneous, components were appropriately bundled

into the miscellaneous drag coefficient approximation.

With the aircraft’s wing airfoil chosen, the calculation of induced drag was

possible. However, a number of assumptions needed to be made. It was assumed that at

cruise, the combination of the air’s viscosity and the local length scales were not enough

to damp out the downwash produced by the canard prior to its interaction with the main

wing. With this assumption made, induced drag of the canard was calculated with an

induced drag prediction code developed in AAE 334, while the main wing’s induced drag

was found assuming the linear summation of the downwash angle of attack produced by

both the canard and main wing. The induced drag prediction found an overall induced

drag coefficient of approximately 0.0175. A technology factor was then applied of 0.93,

which is a result of the advanced technologies included in the design, to yield a final

induced drag coefficient of 0.01002.

The last part of the drag component buildup to be completed was a better

calculation of wave drag. The wave drag coefficient is approximated according to the

following equation.

CD,wave = 20*(M – Mcr)4

Here, Mcr is the critical Mach number, the Mach number at which wave drag first occurs. M cr is

usually considered to occur at a Mach number 0.07 less than the drag divergence Mach number,

MDD. However, the use of a supercritical airfoil pushes MDD forward such that Mcr occurs at a

Mach number only 0.01 less than the MDD calculated for a non-supercritical wing. MDD was

found to be approximately 0.83, yielding a Mcr of 0.82. All of the information needed to

calculate the wave drag coefficient at cruise is now known. Overall, the aircraft was found to

generate a drag coefficient of 0.02665 at cruise. This drag coefficient is less than that created by

the smaller Cessna 172 Skyhawk, but slightly higher than Boeing’s newest premier airliner, the

787.

PERFORMANCEV-n Diagram

33

In the V-n diagram for the aircraft, several factors were considered. First as seen

in Figure 26, the aircraft was limited by the line outlined in red. On the positive load

values, at low velocity (up to 790 ft/s) the aircraft is limited by Clmax. The equation used

for this calculation is listed below. After this velocity, the aircraft is limited by the gust

curve up to the cruise velocity and the first Δn value. This value and the other Δn values

were found by taking the gust speeds from the FAA at their altitude and interpolating for

41,000 ft at each velocity. Finally the curve is limited by the dive velocity at the far end.

This was chosen to be Mach 0.87 for this aircraft. This same process was done for the

negative load factors.

n=

Clmax∗12

∗¿V 2

W 0/S

Figure 26 Aircraft V-n Diagram

Additionally, a payload range diagram was generated for our final design. This

simple plot shows how far the aircraft could fly in any loading scenario. The payload

34

range diagram is shown in Figure 27, and is applicable to a cruise Mach of 0.85 at an

altitude of 41,000 feet. The ranges listed are still air ranges.

0 1000 2000 3000 4000 5000 6000 7000 8000 90000

500

1000

1500

2000

2500

3000

3500

4000

4500Payload Range Diagram

Range (nmi)

Pay

load

Wei

ght (

lbs)

Figure 27: Payload Range Diagram

PROPULSIONThe engine was modeled using a rubber engine approach. The MATLAB code

that models the engine uses the tabular engine data for the CF-34 that was provided in

class. The engine code scales this data based on the sea-level static thrust of the engine

deck and that which is required by the aircraft. The SFC data is also scaled using a

technology factor to account for new technologies such as the use of a geared turbofan.

This technology factor also includes a small correction factor for installation losses. The

sizing code uses the developed engine model to calculate the weight of the fuel required

in each segment of the design mission.

Once the aircraft is sized, the engine and drag models were used to generate thrust

required and thrust available curves for a range of speeds at sea-level and cruise altitude.

35

The thrust available curves are generated by simulating the engine model at full throttle at

a given altitude and a range of velocities and multiplying the resulting thrust by the

number of engines, Figure 28. The thrust required is determined by the drag on the

aircraft. Figure 29 shows that there is a maximum operating airspeed, which corresponds

to a maximum operating Mach number of about 0.89.

100 150 200 250 300 350 400 450 500 5500

0.5

1

1.5

2

x 104

Velocity (kts)

Thru

st (l

bf)

Thrust Required Curve at Sea Level

Thrust RequiredThrust Available

Figure 28: Required and available thrust at sea-level for various airspeeds.

36

250 300 350 400 450 500

2000

2500

3000

3500

4000

4500

Velocity (kts)

Thru

st (l

bf)

Thrust Required Curve at 41000 feet MSL

Thrust RequiredThrust Available

Figure 29: Required and available thrust at cruise altitude for various airspeeds.

STRUCTURESThe placement of major structural components during the preliminary design process was

investigated in order to provide a starting point for future detailed design. Depicted in Figure 30

is a diagram of the major structural components that compose the outer mold line of the aircraft.

Examining historical cutaways from similarly sized and loaded aircraft, two spars were deemed

necessary for the preliminary design process. The upper and lower spars are connected by shear

webbing as indicated on the diagram in green, and the wing and fuselage intersection is filleted

to provide lower aerodynamic interference as shown in blue. Ribs are modeled with equal

spacing to maintain the airfoil’s shape and transfer the distributed skin loads from to the spars.

The vertical tails and canards share a similar construction to the main wings, though with one

main spar in each to account for the lower force distribution and span on each component.

37

Also shown in Figure 30 is a rough sketch of the major components of the fuselage

structure. The fuselage is constructed from structural frames to maintain its cross-sectional

shape, connected with longerons to resist bending and to support the aircraft’s skin. Reinforced

bulkheads are integrated into the main frame at the fore and aft most portion of the cabin to

provide a barrier for the pressurized ends of the main cabin. Sills were modeled above and

below the fuselage cutouts for the doors and windows, and beams with several webbing cross

members were placed on the sides of each door and window to tie the sills together and to

connect them to their respective fuselage frames. The sills and beams are indicated in red and

blue around each door and window in Figure 30.

Figure 30 Overview of Major Structural Members

Load Paths

After modeling a sampling of the major structural components on the aircraft frame, the

major aircraft loads were applied to the structural model in low fidelity to provide evidence of

the structural members needed to carry the specified loads. Distributed lift and weight loads were

applied to the airframe through the skin, and ribs, which was transferred from the ribs to the

wing’s spars, and passed to the fuselage. The drag force on the wings and fuselage are carried by

the aircrafts skin, passed to the ribs, spars, frames, and longerons, and is distributed throughout

the entire aircraft framework along with the fuselage weight. Point loads from the thrust of both

engines are carried through the pylon supports into the beams between the large fuselage frames

38

surrounding the pylons, and is likewise distributed through the fuselage assembly through the

frames and longerons. The major loads applied to the aircraft and a low fidelity model of the

structural members that carry them are shown in Figure 31.

Figure 31 Structural Highlights: Major Loads and Load Paths

Wing-Fuselage Intersection

A more detailed sketch was drawn for the wing-fuselage intersection, and is presented in

Figure 32. The distributed weight, lift, and drag loads carried by the wings are transferred from

the wing’s skin to the ribs, then distributed across the wings spars, and are carried from the spars

to the fuselage. Surrounding the wings and the wing spars at the fuselage are reinforced frames

that are connected with beams above and below the wing intersection, to form a wing box that

accepts the wing cradle formed by the strake assembly. The wing box distributes the loads

imparted by the wing cradle to the longerons, thereby distributing the wing forces across the

entire fuselage frame assembly. Also visible in Figure 32 is the fillet surrounding the root of the

wing at the fuselage intersection, which distributes both the forces carried by the wing skin and

smoothes the aerodynamic transition point between the wing and fuselage.

39

Figure 32 Wing-Fuselage Structural Intersection

Engine Pylons / Mounts

Similar to the wing–fuselage intersection above, a more detailed sketch was also

compiled for the engine pylons and pylon mounts, shown in Figure 33. Carrying the

entire thrust loads of the aircraft, the pylons and pylon mounts distribute a focused large

amount of force over a small cross sectional area. Additional frame reinforcement is

expected in the frames immediately fore and aft the pylon areas. The main engine weight

is supported by a high strength rod spanning the center of the pylon shroud, and bending

and other loads are carried by the pylon shrouds constructed from ribs and a single spar in

a similar manner to the canard and vertical tails. Loads transferred from the pylons and

main pylon support rod encounter a large beam at the fuselage, spanning several

reinforced structural fuselage frames, and the aft cabin pressure bulkhead. The loads

carried by the reinforced frames and bulkhead are then distributed throughout the

fuselage frame assembly by longerons and additional frames.

40

Figure 33: Pylon / Engine Mounting Structural Framework

Landing Gear Integration

Similarly to the engine-pylons and wing–fuselage intersection, a more detailed

sketch was constructed for the landing gear placement in the main wings shown in Figure

34. Investigation of historical landing gear structure for similar aircraft indicated the

landing gear assembly to be supported by beams attached to reinforced ribs outboard on

the main wings. In order to provide adequate wing thickness to accommodate retractable

landing gear, the landing gear must be placed near the twin vertical tails, to allow a large

enough bay into which the landing gear could retract. The loads from the main gear on

landing are transferred through the landing gear assembly, the reinforced ribs, and into

the wing spars to be passed to the fuselage in the same manner as the lift, weight, and

drag forces on the wing.

41

Figure 34: Landing Gear Structural Framework Detail

Material Selection

In order to provide a reduction in empty weight fraction, research was conducted on the

incorporation of a combination of both advanced and traditional materials on the airframe. The

materials selected to be used on the aircraft were chosen based on five main factors.

Strength to Weight Ratio

Fatigue Resistance

Maintenance

Cost

Ice-phobic Characteristics

In order to meet the chosen requirements, the following materials were considered.

Advanced Composites

Advanced composite materials were taken into consideration because of their high

strength to weight ratio and tailoring capabilities. Use of advanced composites on

commercial aircraft has been limited due to material and manufacturing costs, as well as

poorly developed inspection and repair practices. With few exceptions, commercial

airliners have incorporated a maximum structural weight of 15% composites until the last

decade. More recent undertakings such as the Boeing 787 have achieved a composite

42

makeup of up to 50% structural weight, and will help promote the development of more

established inspection and maintenance practices, resulting in potentially easier aircraft

certification.

Our design will incorporate a quasi-isotropic [45°/0°/-45°/90°]s symmetric layup

of AS4-3501-6 carbon-epoxy unidirectional pre-preg on the wing skin and control

surface panels.15 Quasi-isotropic carbon-epoxy layups possess a specific strength of three

times to five times that of the aluminum alloys typically used in wing skins panels.15 Due

to the addition of cooling, ice-phobic coatings, and electrical grounding systems required

in higher composite fraction aircraft based upon historical research, a more modest 50%

reduction in empty weight fraction for the relevant components is expected.

While there are significant potential benefits from the use of advanced composite

materials, a number of difficulties have historically limited their use on aircraft. Current

theoretical failure prediction incorporates a variety of inaccurate assumptions, but this

adversity is becoming largely offset by the growing historical database of composite

aircraft. Additionally, fatigue and fracture detection on composite materials lacks a wide

array of non-destructive techniques, leading to historically difficult certification of

primarily composite aircraft. Recent improvements in ultrasonic scanning, x-ray, and

acoustic emission techniques are projected to increase with the advent of the Boeing 787,

and it is expected that the newly developed field will become sufficient to ease the

certification of primarily composite airframes by the 2020 manufacturing date.

The development of new open-and-closed molding techniques, as well as

injection molded resin transfer methods for cylindrical fuselage sections, have reduced

manufacturing costs for composite aircraft construction. Despite the manufacturing cost

reduction, composite manufacturing costs are still higher than more conventional

manufactured aluminum aircraft. A reduction in the manufacturing cost of composite

materials on aircraft is expected resulting from the combination of a lower parts count

afforded by composite panel layup, as well as mass production advances developed

through the increased use of composites on the Boeing 787.

43

Advanced Aluminum Alloys

While a great deal of research and discussion has been conducted to the use of

composites in the aerospace industry, aluminum alloys have continued to advance as

well. Recent developments in Aluminum-Lithium alloys have been targeted directly at

the aerospace industry due to the high strength to weight ratio, exceptional fatigue

performance, and cryogenic toughness properties. Aluminum-lithium alloys have existed

since the 1950’s, and were developed by adding lithium to aluminum-copper, aluminum-

magnesium, and aluminum-copper-magnesium alloys. Aluminum-lithium alloys promise

superior crack propagation resistance when compared to the more traditionally used 2000

and 7000 aluminum alloys used in airframe construction.

Alcoa’s 2090 series of aluminum-lithium alloys have been used on military

aircraft and the NASA space shuttle to provide significant weight reductions, thereby

allowing heavier payload capacities.16 The increased strength of aluminum-lithium alloys

has promoted the use of less material for the same strength and safety margin compared

even to advanced composites, and are employed up to 20% by structural weight on

commercial aircraft such as the Airbus A350.

Though more fatigue and crack resistant than traditional aluminum alloys, a

number of disadvantages exist for aluminum-lithium alloys in aerospace applications.

Due to the increased strength and crack resistance properties, aluminum-lithium alloys

contain a reduction in ductility and retain a degree of anisotropic properties. Specifically,

some lithium-aluminum alloys exhibit reduced fracture toughness in the transverse

direction from rolling, and promote accelerated fatigue crack extension for structurally

small micro-cracks.16 Despite the listed negative aspects of aluminum-lithium alloys,

their use on aircraft promotes a 1.5 to 3 times increase in expected life, and can be

manufactured, inspected, and repaired using conventional aircraft metal techniques. Due

to their superior strength to weight ratio, aluminum-lithium alloys were afforded the same

technology factor as advanced composite materials for the relative components.

44

Aluminum

Aluminum has been traditionally used in the aviation industry for up to 80% by

weight of typical modern aircraft. Forged and machined aluminum components are

relatively easier and cheaper to manufacture than their more modern material alternatives,

and have decades of reliable data regarding their strength, fatigue, and fracture properties.

Alloys typically used in commercial aircraft include 7075 for high stress applications due

to its high ultimate strength, relatively low weight, ease of machining, anodization

properties, and smooth finish.

Traditional uses of aluminum plate, coils, and sheets will be employed on our

aircraft in the absence of more technologically advanced materials due to their

historically proven performance in harsh environments, reduced manufacturing and

developmental costs, and established testing and repair techniques. It is also expected

that the incorporation of a large percentage of traditional aluminum alloys will provide

easier aircraft certification, and more reliable performance and airframe life data. As

aluminum is the most traditionally used material on modern aircraft, no technology factor

is incorporated for the aluminum aircraft components.

Steel Alloys

Heat-treated steel alloys were chosen for highly loaded components such as the

aircraft’s landing gear. Containing the most stringent performance requirements, the

landing gear is subject to severe loading conditions in a variety of environments.

Today’s commercial and military aircraft use 300M, HP9-4-30, and newer AF-140 or

AerMet 100 steel alloys. Additionally, the use of Ferrium S53 high strength stainless

steel alloys have been proposed due to strength to weight ratio improvements and

corrosion resistance over more traditional steel alloys.

Latrobe Specialty Steel has recently licensed the production of Ferrium S53 as a

high strength, high toughness, and corrosion resistant alloy suitable for aircraft landing

gear. Ferrium S53 has also been labeled an environmentally friendly landing gear by the

U.S. Air Force, having met the U.S. Air Forces requirement of a landing gear material not

requiring a toxic cadmium coating.17 Latrobe’s S53 steel is expected to expand into the

45

commercial aviation market in the foreseeable future, with other non-landing gear

aerospace applications such as jet engine bearings on the near horizon.17

While a significant empty weight fraction reduction from the use of Ferrium S53

is not expected on our aircraft application, it is an attractive material for the landing gear

components due to its superior toughness, corrosion resistance, and environmental

friendly manufacturing techniques through the lack of toxic coatings. Additionally,

currently manufactured Ferrium S53 landing gear provides a manufacturing cost

reduction to titanium alloys, and is expected to decrease further in price as their use

becomes more widespread.

Material Choice Benefits

The use of exotic materials on aircraft promises a significant reduction in empty

weight fraction, however one must take into account the costs associated with each

material, including developmental, manufacturing, and maintenance costs, ease of

certification, and additional system weight requirements. We expect a more conservative

structural weight reduction applied as a technology factor of 20% from the use of

advanced composite materials and Lithium Aluminum alloys on the aircraft’s skin panels,

nose, and leading and trailing edge surfaces. The material choices for major aircraft

components are listed below in Table 9.

Table 9: Aircraft Component Material Choices

Component Material Advantages

Fuselage skins and wing stringers Aluminum Alloys (Al-Li)

Better Fatigue Crack Growth (FCG) performance reduces

structural weightCanard, Control surfaces

and wing skin panelsCarbon-Epoxy Composite

LaminatesResistant to damage at high

temperaturesNose, Leading and Trailing

edgesCarbon-Epoxy Composite

LaminatesHigh fracture toughness and

yield strength

Landing Gear Steel Alloy (Ferrium S-53) High strength, corrosion resistant

46

47

WEIGHTS AND BALANCEAs discussed in past reports, the empty weight for the aircraft was calculated on a component

weight buildup method using equations from the Transport Aircraft category of Daniel Raymer’s

design textbook.8 For our final and best configuration; this led to the weights table listed below

as Table 10.

Table 10: Component Weight Table

Component Weight (lbs) Component Weight (lbs)

Gross Weight 71,331 Fuel System 308

Empty Weight 38,079 Flight Controls 683

Fuel Weight 31,481 Installed APU 440

Wing 6,913 Instruments 262

Canard 598 Hydraulics 180

Vertical Tail 1,085 Electrical 943

Fuselage 7,224 Avionics 471

Main Landing Gear 873 Furnishings 3,054

Nose Landing Gear 103 Air-conditioning 199

Nacelle 1,628 Anti-ice 139

Engine Controls 90 Crew 800

Engine Starter 130 Payload 3,960

CG Travel Diagram

Figure 35 below shows the location of the aircraft’s center of gravity for various

loading conditions. Four points in particular are listed in Table 11 and their

corresponding locations are shown on the plot. Point A represents the maximum takeoff

weight. Flying the design mission corresponds to a movement from point A to B, and the

various points between A and B on the plot correspond to the endpoints of each mission

leg. Moving from points B to C corresponds to the unloading of all passengers and crew.

Moving from points C to D corresponds to the refueling of the aircraft, and moving from

points D to A corresponds to the loading of all passengers and crew to full capacity.

Over the course of aircraft operation, the center gravity travels in approximately an eight

percent range of the total fuselage length.

48

62 63 64 65 66 67 68 69 70 71 7230,00035,00040,00045,00050,00055,00060,00065,00070,00075,000

CG Travel Diagram

CG (% Fuselage Length)

Tota

l Wei

ght (

lbs)

Figure 35: Aircraft Center of Gravity For Varied Loading Conditions

Table 11: Four Loading Conditions and Center of Gravity Points

Point Number of Passenger

s

Number of

Crew

Fuel (lbs)

A 16 4 37,750

B 16 4 0

C 0 0 0

D 0 0 37,750

STABILITY AND WEIGHTSControl Sizing

Crosswinds

Quite often a runway is oriented in such a way that landings must be made while the

wind is blowing across the runway rather than parallel to it. When this is the case, additional

complexities are introduced into the approach and landing process. If left uncorrected, the

crosswind acting upon the aircraft will continually cause the plane to drift in the direction of the

wind, making a conventional landing impossible. One method for counteracting the effects of the

49

D

CB

A

crosswind is the crab method. In this method, the pilot is required to yaw the aircraft into the

wind at such an angle that the flight path remains aligned with the centerline of the runway, as

seen in Figure 36. However, to ensure the pilot will have adequate control to prevent the continual

drifting and handle crosswind landings, special considerations must be taken with regards to the

vertical tail size.

Figure 36: Crosswind Landing

In order to maintain a steady heading into the wind, the moment generated by the

crosswinds must be opposed by a moment of equal magnitude generated by the deflection of the

rudders on the vertical tail. The force each vertical tail can generate is a function of its coefficient

of lift, its distance from the center of gravity, and its planform area. If everything other than the

planform area is considered fixed, the resulting vertical tail area is the minimum size the vertical

tail must be to adequately perform under crosswind conditions.

One Engine Inoperative

In addition to crosswinds, the one engine inoperative case can potentially place large

constraints on an aircraft’s minimum vertical tail size. The one engine inoperative case is

characterized by the loss of one engine during flight which results in asymmetrical thrust. This

asymmetrical thrust generates a yawing moment that must be balanced by an opposing force

created by deflecting the rudders on the vertical tail, as seen in Figure 37.

50

Figure 37: One Engine Inoperable

Much like the calculations for crosswinds, the necessary force each vertical tail needs to

generate an equal opposing moment will determine the minimum area needed for the vertical

tails. Having sufficient vertical tail area will allow the aircraft to maintain a constant heading

angle with only one engine operable.

For this aircraft specifically, the one engine inoperable condition is less constraining than

the crosswinds condition with respect to vertical tail sizing. The primary reason that the one

engine out condition is less constraining is because the engines are mounted directly onto the

fuselage near the centerline of the aircraft. Therefore, when one of the engine stops generating

thrust, the resulting moment from the operable engine is still relatively small due to its small

moment arm. With a smaller moment, a smaller amount of force from each vertical tail is

required, and ultimately smaller vertical tails are needed. By using the vertical tail size as

determined by the crosswind condition, the airplane will be designed to withstand both adverse

flight conditions.

Horizontal Stabilizer Sizing

In the specific case of the canard design, the canard is the horizontal stabilizing

surface. An equation from Raymer, shown below, was used as a preliminary sizing

condition since the horizontal stabilizer was also a lifting canard and thus required a

larger surface to provide lift.8 The final size was calculated by multiple iterations through

the sizing code to find an ideal surface area.

51

• CHT-Tail Volume Coefficient (1.00 for Jet Transport)

• LHT-moment arm (quarter chord of wing to quarter chord of tail)

• SHT-surface area of horizontal tail

• C-main wing chord

• Sw-surface area of the main wing

Control Surface Sizing

The aileron, rudder and elevator control surfaces were sized based on historical

data provided in Raymer. This data is presented in Figure 38, and Table 12, shown

below.8 By initially calculating the surface size of the wing, horizontal stabilizer, and

vertical stabilizers, the ailerons, elevators, and rudders could be sized respectively.

Figure 38: Raymer Figure Number 6.3

52

Table 12: Raymer Table Number 6.5

Static Margin

Finally, it was important to predict the stability of the aircraft by determining its

static margin. This calculation required two values, the location of the center of gravity

and the location of the aerodynamic center. The center of gravity was calculated using

the weights and locations of each major piece of the aircraft. The aerodynamic center

was less straightforward. For conventional configurations, the aerodynamic center can

usually be approximated at the quarter chord with acceptable accuracy. For an aircraft

with a lifting canard, the aerodynamic center will undoubtedly shift forward. Since no in-

depth aerodynamic evaluation could be performed, it was assumed that the aerodynamic

center of the canard configuration moved from the quarter chord of the main wing to the

leading edge of the main wing. With this assumption, the center of gravity was

calculated and the static margin was found to be 9.54% of the main wing chord.

Research shows that many transport aircraft have a static margin of 5-10%. Therefore

our predicted static margin value of 9.54% was within the normal range. It was also

noticed that the center of gravity and static margin varied widely with small changes in

the location of the engines. This indicates that a more thorough analysis of engine

placement would be required during the preliminary design phase, in order to ensure

acceptable stability characteristics.

53

NOISE 1 One of the main N+2 goals is to significantly reduce noise. Because of that,

Team 2 predicted the noise of our aircraft using methods and equations commonly used

in acoustic engineering. In order to carry out simple noise estimation, a few assumptions

were necessary. First, it was assumed that the primary noise source for the aircraft was

the engine, and that the engine noise came from the fan and the jet exhaust. Additional

information about the engine needed to be approximated, such as the volumetric flow rate

through the core, the exhaust temperature, and the exhaust velocity. All necessary

quantities were approximated based on research of similar engines and on simple cycle

analysis results. After establishing these assumptions and approximations, the following

procedure was used to predict the noise of the aircraft.

Step 1 – Establish the power of each noise source

In this step, the power of both the engine fan and exhaust jet were calculated. The

jet sound power can be calculated using the equation listed below.

P jet=ε M 5 ρoV 3 A

2[Watts ]

In the above equation, ε is radiation efficiency, A is nozzle area, ρo is density of

ambient air, M is Mach number, and V is flow velocity. Since the radiation efficiency is

difficult to predict or approximate, an approximation for the entire term εM5 can be found

from Figure 39 below.

1 Unless explicitly stated otherwise with an exponent reference number, every equation, chart, graph, and table listed in the Noise section comes from reference number 10.

54

Figure 39: Radiation Efficiency Factor vs. Mach Number

Obtaining the sound power of the fan is not necessary, as the noise estimation

process for fans starts with the sound power level rather than just the sound power.

Step 2 – Obtain the sound power level of each source

For the jet, the power value calculated in step one can be converted to a sound

power level with the following equation. This equation uses the standard acoustic

reference power level of 1x10-12 watts and accounts for the difference in temperature

between the ambient air, Ta, and the exhaust temperature, T.

Lw=SPL=10 log(P jet

10−12 )+20 log( TT a

)[dB]

The sound power level of the fan can be predicted with a slightly more

complicated scheme. First, the type of fan must be selected and the corresponding octave

band sound power levels must be selected from Table 13 listed below.

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Table 13: Band Sound Power for Varied Fan Types

A correction factor must then be calculated and added to all values taken from the

above table. The correction factor, Δ, can be calculated according to the equation below,

where Q is volumetric flow rate in cubic meters per second and p is the total pressure in

kPa.

∆=10 log Q+20 log p [dB]

Notice that the last column of Table 13 is titled “BFI,” which stands for Blade

Frequency Increment. This value is measured in decibels and must be added to the

octave band level in which the frequency of the rotating blades lies. So after pulling the

appropriate data from the table, then adding the correction factor to all octave bands

based on volumetric flow rate and pressure, step three in this process is to add the BFI

56

value to the single octave band that contains the rotating blade frequency. This frequency

can be calculated according to the following equation, where n is fan speed in revolutions

per minute and N is the number of blades.

f =nN60

[Hz ]

The values obtained by summing steps one to three are the total sound power of

the fan. In order to obtain the sound power level radiated from the inlet or outlet, the

final step in the process is to subtract 3 dB. After completing all four steps, the radiated

sound power level of the fan is found for each octave band. The individual octave bands

can be summed into a single sound power level according to the following equation.

LW , fan=SWL fan=10 log(Σ10Loctave /10)

Step 3 – Calculate the sound pressure level from the sound power level and distance from source

With the sound power levels of the fan and the jet known, the sound pressure

level can be calculated based on the distance from the source. This calculation also

assumed spherical wave propagation and includes the reflected wave from the ground

since the noise sensors would likely be placed near the ground. The step also includes an

estimation of the aircraft’s altitude 6000m after takeoff and 2000m prior to landing.

These values were estimated using the sizing code and standard approach requirements.

With all of the necessary information acquired, the sound pressure level for the fan was

calculated with the following equation where r is the distance from the source, and will

vary depending on the three standard measurement locations.

Lp , fan=SPLfan=Lw ,fan−20 logr−10 log 4 π+3[dB ]

Additionally, the factor of 4π is a result of the spherical wave propagation assumption, and the

+3 dB correction accounts for the reflected wave.

The equation for the exhaust jet is similar but includes another factor called the

directivity index, which predicts a difference in sound pressure level based on the sensor’s

57

angular position to the exhaust jet. The sound pressure level of the jet can be calculated from the

equation below, and the directivity index, DI, can be found from Figure 40.

Lp , jet=SPL jet=Lw , jet+DI−20 log r−10 log 4 π+3 [dB]

Figure 40: Sound Directivity Index vs. Angle from Jet Axis

Step 4 – Apply A-weighted noise correction

With the sound pressure level of each source known, the decibel values must be

adjusted to better represent the response of the human ear. A-weighted adjustments were

used, which require that the dominant frequency be known. The fan frequency has

already been calculated, and the jet frequency can found according to the following

equation where α is the Strouhal number and assumed to be 0.15, V is the velocity, and D

is the jet diameter.

f =αVD

[ Hz]

58

A-weighted adjustments can be made with the standard acoustic engineering

adjustments provided in Table 14.

Table 14: Standard Acoustic Engineering Adjustments

Step 5 – Calculate loudness measurement

Using a plot of equal loudness contours, the loudness of each source can be found

based on its sound pressure level in decibels and dominant frequency in hertz.11 This is

done quite simply by using the plot shown in Figure 41, with the decibel level along the

Y-axis and the frequency along the X-axis, to find the loudness in Noy of each source

from the approximate location between loudness curves.

59

Figure 41: Band Sound Pressure vs. Frequency

It is at this step in the process that the two sources, fan and jet, can be

arithmetically summed to obtain a single numerical value representing the engine as a

whole. Loudness values in Noy can be accurately summed to combine sources.

Step 6 – Convert loudness values to Effective Perceived Noise Level

Finally, with the engine noise represented as a single loudness value measured in

Noy, the Effective Perceived Noise Level (EPNL) can be found according to the

following equation.12

EPNL=33.3 log(Noy)+40+10 log( t20 )+F

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Here, Noy is simply the loudness value found in the previous step, t is the time in seconds at

which the sensor is exposed to the noise source within 10 dB of its maximum sound pressure

level, and F is a correction factor for pure tones, which is typically found to be +3 dB. The peak

noise level of an aircraft flyby usually occurs over a very small time interval, so t used in

calculation can be approximated by any value between 1 and 20 seconds. In this specific case,

20 seconds was used in order to maximize the predicted sound level and obtain a conservative

estimate.

Following the process outlined above, the noise of each engine considered was

found. The results of this analysis are provided in Table 15, which shows the EPNL

noise estimation of the geared turbofan and unducted fan. From this comparison, it is

clear that the geared turbofan is considerably quieter than the unducted fan. This was one

of the primary reasons why the geared turbofan was selected over the unducted fan.

Table 15: Engine Noise for Various Measurement Locations

Geared

Turbofan

(EPNL dB)

Unducted

Fan

(EPNL

dB)

Sideline 97 102

Takeoff 90 95

Approac

h

97 100

It is important to note that the above table lists the EPNL decibel levels for the

uninstalled engine. This does not include the sound shielding devices used on our aircraft

of the outboard vertical tails and low wing. Because of this, it was important to predict

the noise of the installed engines. This was done by following the procedure outlined

above, but subtracting 15 dB from the sound pressure levels calculated in step three. The

subtraction of 15 dB was chosen because in acoustic engineering, the noise control

61

method of placing a barrier between the source and the sensor can be approximated by a

drop in the sound pressure level of 10 to 20 dB. The exact value would depend on the

material of the barrier, the size of the barrier, and the distance from the source. The

median value of 15 dB was chosen, and applying this correction yielded the installed

noise estimation of the aircraft presented in

below.

Table 16: Corrected Airplane Noise Estimate

Airplane

Noise (EPNL

dB)

Sideline 87

Takeoff 80

Approac

h87

Total 254

COST PREDICTIONAcquisition Cost

The cost estimating relationships that were employed to predict the development,

manufacturing, and certification costs of this system were part of the Development and

Procurement Costs of Aircraft model (DAPCA IV). This model was developed by the RAND

Corporation for a variety of aircraft types, and is discussed in Daniel Raymer’s Aircraft Design:

A Conceptual Approach.8 The model consists of a series of relationships that were determined

from a statistical analysis of previous program costs. The DAPCA IV model predicts the

aircraft’s research, development, testing, and evaluation costs (RDT&E); as well as the

“flyaway” cost, which includes airframe production, and the integration of engines and avionics.

The costs output by the model were in constant 1999 dollars, and were converted to 2009 dollars

using the U.S. consumer price index. To apply the relationships in the DAPCA IV model, seven

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aircraft specification inputs were required: empty weight, maximum velocity, production

quantity, number of flight test aircraft, total number of engines needed for fleet, the cost of each

engine, and the cost of avionics. The model uses these inputs to predict the number of labor

hours required for engineering, tooling, manufacturing, and quality control; and the direct costs

associated with development support, flight test, and manufacturing materials. From these

predicted values, the total cost of developing and producing the aircraft was estimated by taking

into account the costs of labor.

The values used as the inputs for the cost model were determined using the team’s design

specifications, data from previous development programs, and market forecasts. The aircraft’s

empty weight (38,100 lbs.) and required number of engines (2) were readily available details of

the team’s concept. The maximum velocity was approximated as 527 knots based on values from

similar aircraft. In addition, since the number of flight test aircraft typically used in past

programs ranged from two to six, three was chosen as a conservative number for a small volume

production run.8 The cost of avionics was estimated by assuming a price of approximately

$3,000/lb and using the estimated weight of the avionics on board (471 lbs); the cost per pound

was based on typical values supplied in Raymer.8 The number of aircraft to be produced was

estimated using the business aircraft market forecast supplied by Honeywell.2 This forecast

predicted a relatively stable demand for ultra-long range (ULR) business jets with moderate

growth in the high-speed ULR segment. With these market conditions in mind, the production

runs of the Gulfstream G550 and Bombardier Global Express family were studied since these

aircraft were within the same category and had been in production for more than 5 years. These

production runs were in the neighborhood of 180 to 250 airframes. Therefore, it was decided to

use 150 airframes as a conservative value for cost estimation. This value accounted for the

projected stability of the market, as well as the impact of being a new entrant.

The cost of each engine was approximated using a separate cost estimating

relationship (CER) originally developed for turbojet engines.8 This CER approximated

the production cost of an engine from its maximum thrust, maximum operating Mach,

and turbine inlet temperature. The maximum thrust used was the value that had been

determined by the team’s design (11,900 lbs). The maximum operating Mach for the

engine was approximated as Mach 0.95, and the turbine inlet temperature was

approximated as 2600⁰ Rankine; which was based on typical values for modern

63

turbofans. In addition, a correction factor of 1.2 was used to compensate for the increased

cost of production of a Geared Turbofan (GTF) relative to a turbojet. This resulted in an

estimated engine production cost of $3.6 million per engine (2009 dollars).

Using all of these specifications, numbers, and costs as inputs for the model, the

result was an estimated purchase price of $49.7 million per aircraft with a total program

cost of $6.71 billion (2009 dollars). Aside from summing all of the costs included in the

model, an investment cost factor of 1.1 was also used to account for the cost of money

and the team’s profit.8 The projected price of the design was compared with comparable

aircraft currently on the market, and appeared to be a reasonable estimate. The

Gulfstream G550 and Bombardier Global Express XRS had new list of prices of $49

million and $48 million, respectively.6

Operating Costs

After an approximate purchase price had been determined, the operating cost of

the aircraft was estimated. The cost of operating the aircraft was broken down into fuel

costs, crew salaries, maintenance labor, maintenance materials, insurance, and

depreciation. Each of these costs were estimated separately and then used to determine an

overall cost per flight hour.

Fuel costs were estimated from the aircraft’s expected fuel flow and utilization.

An approximate fuel flow of 2,300 pounds per hour was determined using the aircraft’s

design mission, predicted fuel weight, and a correction factor for off-design performance.

The utilization of the aircraft was then approximated as being 500 flight hours per year

with 200 cycles. This was the lower bound of the range of flight hours provided in

Raymer, and was thus expected to yield conservative hourly cost figures.8 In addition, the

number of cycles used implied an average flight duration of 2.5 hours, since the typical

operating mission of the aircraft was expected to be domestic. Next, the price of jet fuel

was considered. After reviewing the current price of Jet A at numerous airports, a price of

$4.50 per gallon was chosen for the purposes of cost estimation. The resulting yearly fuel

cost was $755,000, assuming 500 flight hours per year, or $1,510 hourly.

Crew salaries were approximated from a cost estimating relationship developed

using data from Boeing.1 This relationship estimated crew cost per block hour for a three

64

person crew in 1999 dollars. The annual number of block hours for the aircraft was

determined by assuming that an additional 20 minutes would be spent on each flight

taxiing and complying with air traffic control procedures. The crew cost per block hour

was then estimated from the appropriate CER, which took into account the aircraft’s

cruise velocity (487 knots at altitude) and gross takeoff weight (71,300 lbs). The resulting

crew cost was $537 per block hour, or $268,000 a year (2009 dollars).

Maintenance expenses were estimated using the expected maintenance

requirements of the aircraft, the cost of labor, and CERs. All maintenance to be

performed on the aircraft was included in an average figure of Maintenance Man hours

per Flight Hour (MMH/FH). An MMH/FH of 3 was assumed for this design, which was

in the range considered typical for a business jet.8 The number of maintenance man hours

per year was then estimated from the MMH/FH and the expected number of flight hours

(500). The cost of a maintenance man hour was then assumed to be approximately equal

to the hourly manufacturing wrap rate. This resulted in an estimated maintenance labor

cost of $282 per flight hour. The cost of maintenance materials was then estimated using

CERs for civilian aircraft provided in Raymer from RAND.8 These CERs predicted the

material cost per flight hour and per cycle using the estimated cost of the airframe, the

cost of the engines, and the number of engines on each aircraft. The result was an

estimated maintenance materials cost of $619 per flight hour.

Depreciation and insurance costs for the design were estimated from data

gathered on similar aircraft models. Depreciation was approximated as a linear schedule.

After reviewing the asking prices of numerous ULR business jets on the used market, a

depreciation rate of 10% per year was selected. This rate would be close to the amount

expected within the first 5 years of ownership, but would be considered a fairly

aggressive estimate after that. Insurance was estimated assuming that the aircraft would

be insured for its full hull value, at a rate of 0.32 percent. This rate was taken from a cost

evaluation of the Gulfstream G550 performed by Conklin & de Decker.9

Adding together the estimated costs of fuel, crew, maintenance, insurance, and

depreciation, an operating cost of $8,500 per flight hour was calculated (2009 dollars).

Without depreciation, the predicted cost was $3,400 per flight hour. For comparison, a

detailed cost analysis of a Gulfstream G550 performed by Conklin & de Decker predicted

65

a total cost of $5,500 per flight hour excluding depreciation, with an assumed utilization

of 400 hours per year.9 Although this is significantly more expensive, the estimates for

the G550 used an average fuel burn of 3,300 pounds per hour; which is significantly

higher than the fuel burn projected for this team’s concept. In addition, the cost analysis

completed here ignored the impact of hangar rent, landing fees, catering, and crew

training on operating costs.

66

SUMMARYThe final concept of our designed aircraft can be seen in Figure 42. The concept is

a high speed, long range aircraft which is also environmentally friendly. The aircraft has

lifting canards, two vertical tails, geared turbo fan engines and spiroid wingtips.

Figure 42: Final Concept Design

The requirement compliance matrix associated with the aircraft is shown in Table

17, and was used to verify that all the requirements of our customers were met.

Table 17: Requirements Compliance Matrix

Performance Characteristics Target Threshold Current

67

Still Air Range 7100 nm 6960 nm 7100 nm

MTOW Takeoff Ground Roll 4000 ft 5000 ft 3900 ft

Max. Passengers 16 8 16

Volume per Passenger per Hour (Design)

13.3 ft3/(pax⋅hr)

2.28 ft3/(pax⋅hr) 20.7 ft3/(pax⋅hr)

Cruise Mach 0.85 0.8 0.85

Initial Cruise Altitude 41000 ft 40000 ft 41000 ft

Cumulative Certification Noise Limits

274 dB 274 dB 254 dB

Cruise Specific Range 0.3 nm/lb 0.26 nm/lb 0.31 nm/lb

Loading Door Sill Height 4 ft 5 ft 5 ft

Operating Cost $4100/hr $4300/hr $3400/hr

Based on Table 17, the majority of the requirements in the requirement compliance

matrix were met with the only one requirement not meeting the target value, which was the

loading door sill height. However, for this height, the target value of 5 feet was met.

In addition to the above requirements, it was also important to our team to meet

the NASA N+2 goals if possible. As seen in the table below, none of the NASA N+2

goals were completely met. The NASA N+2 goals were difficult to meet due to their

overly ambitious targets. In order to meet these goals, more development would be

needed for engine performance, aerodynamic design, and materials used.

Table 18: Compliance with NASA N+2 Goals

Criteria Goal Our Aircraft Achieved

68

Noise -42 dB below Stage 4 -20 dB No

Emissions -75% -50% No

Fuel Burn -40% -25% No

Takeoff Field Length -50% -33% No

Further work would be required to develop the maturity of this concept to a level that is

acceptable for the preliminary design phase. This work would include more detailed

aerodynamic and structural analyses, and a more thorough approach to predicting engine

performance. The aerodynamic analysis that was performed for our aircraft was predominantly

conceptual and theoretical, and more experimental data would be helpful. More in depth detail

for a final design would be needed, using tools such as computational fluid dynamics tools, and

wind tunnel testing. The engine performance could also be more detailed through further analysis

and a more accurate engine deck could be built. Further analysis on the structural design, its

load paths, and the materials used needs to be incorporated. Sizing of the spars and ribs along

with fatigue and temperature analyses would need to be done using finite element modeling to

evaluate the stresses and strain present in the aircraft. With this additional work, the initial

design phase could be completed.

69

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