EXECUTIVE SUMMARYTeam Two is designing a business jet aircraft that is both environmentally responsible

and meets the needs of a selected target customer. These two major drivers for the design of the

aircraft were examined in the previous System Requirement Review document which yielded the

main requirements of a long range, high cruise speed, and low environmental noise and

emissions aircraft.

Many possible designs and advanced technologies were analyzed in order to develop an

aircraft that best meets these requirements. The initial aircraft designs considered included a

hybrid wing aircraft, several conventional aircraft with different engine placements and tail

configurations, and canard aircraft with aft mounted wings and various engine and vertical tail

placements. These aircraft were evaluated and compared until the two best were ultimately

selected. One of these remaining aircraft was a conventional aircraft similar to a G550 except

using an advanced engine and some composite materials. The second was a canard aircraft with

an aft mounted swept back wing using the same advanced engines and composite materials.

In order to meet the NASA N+2 environmental goals, drastic changes will have to be

made in comparison to modern business jet aircraft. To do this, advanced technologies were

evaluated for their potential use in 2020. Two new engine types that are currently in research

phases were examined: an unducted turbofan and a geared turbofan. Using one of these two

engines, and the addition of light weight materials, may provide enough benefit to meet the

goals.

Current calculations show that the conventional aircraft with the advanced technologies

will weigh about 71,000 lbs and the second design using canards and an aft wing, also with the

same advanced technologies will weigh 73,000 lbs. Both aircraft have been determined to have

adequate static stability. Further analysis will be done to determine more precisely which of the

two aircraft designs and which of the engines will best provide an environmentally responsible

aircraft to the customer.

Table of contents

MISSION STATEMENT............................................................................................................................1

DESIGN AND OPERATING MISSION....................................................................................................1

AIRCRAFT CONCEPT SELECTION........................................................................................................2

ADVANCED TECHNOLOGIES..............................................................................................................13

SIZING CODE..........................................................................................................................................16

ENGINE SELECTION..............................................................................................................................30

SUMMARY..............................................................................................................................................39

NEXT STEPS............................................................................................................................................39

REFERENCES..........................................................................................................................................41

APPENDIX A...........................................................................................................................................42

APPENDIX B...........................................................................................................................................46

APPENDIX C...........................................................................................................................................50

MISSION STATEMENT

It is the mission of Team 2 to be the primary systems integrator of a cost effective, long

range, environmentally friendly business aircraft that is able to serve as a high speed

transportation solution for its customers. The design will have a reduced impact on the

environment by meeting NASA’s N+2 criteria while still remaining competitive in the ultra-long

range business jet market.

DESIGN AND OPERATING 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. The design mission consists of

eight mission legs between nine points as illustrated in the following figure.

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. From points 1 to 2 is the climb portion of the mission where the aircraft climbs at best

rate to an altitude of 42,000 feet. From there the aircraft enters the cruise leg of the

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mission, between points 2 and 3, and begins cruising at a Mach number of 0.85 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, climb to an altitude of 5000 feet at

best rate climb, points 5 to 6, and commence cruise to an alternate airport 200 nautical

miles away. Once 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 when it

comes to a stop.

While the design mission is the optimal, most efficient use of this aircraft, several

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

would be flying from New York to Van Nuys. The distance between the two cities, which is

2146 nautical miles, falls well within the maximum still-air range of 6350 miles. To

compensate for the largely unused range, the aircraft can then be flown at its maximum

Mach number of 0.9 at a maximum capacity of 16 passengers. This range tradeoff allows for

tremendous flexibility in speed and capacity for shorter ranged flights.

AIRCRAFT CONCEPT SELECTION

Following the system requirements phase of design, concept generation and selection was

required to further the design process. After great thought and consideration, the team was able

to generate eight initial designs, as shown in Figure 2, for further analysis. Rather than

subjectively choosing one of the eight concept designs, the team implemented Pugh’s method as

a means to add objectivity to the decision making process. The first step was to establish a set of

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criteria against which all of the proposed aircraft would be evaluated. The house of quality that

was constructed during the design requirements phase helped in the establishment of evaluation

criteria. This provided criteria that were centered on the needs of the customer and not the

desires of the designers. Ultimately, ten different criteria, as listed in Table 1, were used in the

first iteration of Pugh’s method.

Figure 2: Eight Proposed Aircraft Designs

The next step of Pugh’s method was the creation of a matrix relating the established

criteria to each of the proposed designs. The matrix’s purpose was to help clarify how each

aircraft concept rated under each criterion relative to a benchmark aircraft. Based on the selected

benchmark aircraft, all other concepts were evaluated on whether or not they were better, worse,

or equivalent to the benchmark aircraft for a given criterion. These relationships were illustrated

in the matrix using the symbols ‘+’, ‘-’, and ‘s’ respectively. Each decision was made with

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cursory calculations whenever possible, relying on the team’s best collective engineering

judgment when not possible.

After evaluating each aircraft in the matrix against every specified criterion, the total

number of positive, negative, and equivalent marks were separately compiled in order to

maintain evaluation details. For instance, if positives and negatives were allowed to cancel each

other out, it would be unclear whether or not an aircraft received a score of zero due to average

qualities or because all of its positive aspects were offset by negative ones. Evaluating the

positives and negatives allowed the team to clearly define the favorable and unfavorable aspects

of each individual aircraft. The matrix comparing the eight initial designs to the ten criteria is

shown in Table 1.

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Table 1: Concept Selection, Pugh's Method 1st Iteration

1 2 3 4 5 6 7 8

Criteria Baseline

Cruise Drag + + s s + + + +

Weight - s s s s - - -

Ability to accommodate UDF - s - s s s s s

Cabin Noise s s - s s s s s

Environmental Noise + s s s s s + +

Landing Gear s s - s s s s +

Window Placement + + - s + s + -

Attractiveness s + - s - + + +

Pressurization s s s s s s s -

Static Margin - - s s - + - -

total S's 4 6 5 10 6 6 4 2

total +'s 3 3 0 0 2 3 4 4

total -'s 3 1 5 0 2 1 2 4

Design Number

After studying the different ratings given to each aircraft, the team decided to repeat

Pugh’s method using the top four aircraft designs determined by the first iteration. This was done

because the first iteration of Pugh’s method did not define a single best design for the given

mission, but rather judged each aircraft qualitatively against the chosen benchmark. The

benchmark aircraft clearly cannot receive a positive or negative rating for any criterion, since it

is equivalent to itself. Thus, two iterations of the method were completed, each with a different

benchmark aircraft.

Examining the outcomes of the first Pugh’s method iteration, it was decided that the

canard equipped aircraft could offer a potential advantage in cruise drag over conventional tube

and wing designs. The overall weight of each design was compared by counting the number of

control surfaces used and observing the general structural complexity of the configuration. No

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material considerations were incorporated into these decisions as they could apply to any of the

designs. An increased number of surfaces, or additional structural requirements such as those

found on a blended wing body, were deemed less favorable for overall weight than a

conventional aircraft configuration of the same size.

In order to meet the project’s goal of having a reduced environmental impact, it was

desired that the airframe be able to utilize a variety of high speed engine technologies, including

unducted fans. To this end, aircraft with rear pylon mounted engines similar to the baseline

model were assigned equivalences to the baseline, whereas designs that would place the engines

in locations interfering with the cabin from vibration, noise, and safety perspectives were

assigned negative ratings. Specifically relating cabin noise relative to the baseline design, a

negative rating was assigned to engine locations in line with the center of the main cabin area.

Further in regard to addressing environmental impact, the environmental noise of each

proposed design was considered. As every design was evaluated with the possibility of using a

variety of engine types, the aircraft structure was the primary consideration for this criterion.

Designs providing noise shielding due to wing placement, vertical stabilizer placement, or

having shielding due to a blended wing configuration were all given positive ratings relative to

the baseline. The placement of each design’s landing gear was also considered. Designs

requiring excessively long and robust landing gear were given a negative rating, whereas the

blended wing body design with a structural framework capable of mounting the landing gear in a

variety of locations was assigned a positive rating relative to the baseline.

Window placement and an unobstructed cabin view were also evaluated for each design,

and aircraft with aft-mounted wings provide a clear window view, which was deemed superior to

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the baseline configuration. Difficulties associated with placing an abundant number of windows

in a blended wing body configuration resulted in a negative rating for the design relative to the

baseline. Though a subjective and hotly debated criterion, each design’s attractiveness and

visual appeal was also considered. By majority vote, negative values were assigned to the high

wing design, as well as aircraft with less sleek lines. The team agreed that canard designs with

aft wing sweep and more futuristic configurations were more visually appealing than the baseline

configuration. In addition, it was decided that the forward wing sweep included in design one

made it less visually appealing.

Each of the eight designs were investigated in relation to two remaining criteria, the first

of which was the complexity of cabin pressurization. Every design with the exception of the

blended wing body contained a relatively easy to pressurize cylindrical fuselage and these

designs, therefore, received equivalency ratings. The blended wing body was assigned a

negative value relative to the baseline due to its non-cylindrical shape. Finally, static margin and

stability were considered for each design. It was decided that canard aircraft, with the exception

of the tri-surface aircraft, would provide a smaller static margin and contain potential center of

gravity challenges, which earned these designs a negative value relative to the baseline. The

limited historical data available suggests center of gravity difficulties and a small static margin

on existing blended wing body designs, thereby also providing it a negative value relative to the

base design. The tri-surface aircraft was given a positive value in relation to static margin, as

data available for the minimalistic number of tri-surface aircraft suggested an ability to design

for a larger static margin.

After evaluating each of the eight designs relative to the initial ten criteria, design

numbers two, four, six, and seven were chosen as the best designs and were used in the next

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iteration of the method. Out of these four designs, number six was chosen as the new benchmark.

Design six was chosen because it was very dissimilar to the remaining aircraft. It was important

to choose a new baseline model so that the outcomes from the second iteration would not be

identical to the first, and would allow the previous baseline, design four, to be rated.

It was noted during the first iteration of Pugh’s method that there were no criteria related

to development cost. In order to improve the value of the second iteration, development cost was

added to the matrix for the four remaining designs, as shown in Table 2. Each design was then

reevaluated for every criterion based on the new benchmark aircraft. Ratings were assigned in a

manner similar to the first iteration, with the same considerations taken into account.

Table 2: Concept Selection, Pugh's Method 2nd Iteration

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Using the results of the second Pugh’s method iteration, two design choices were selected

as best meeting the evaluation criteria and requiring further analysis. Designs four and seven had

the fewest number of negative ratings, with design seven having the most positive ratings during

the second iteration. Since the goal was only to narrow the field of designs down to two options

for additional analysis, further iteration was avoided and designs four and seven were selected.

Designs four and seven, selected by the Pugh’s method process, are henceforth referred to as

concepts one and two, respectively, and are shown in Figure 3.

Figure 3: Concept One, Formerly Design Four (left) and Concept Two, Formerly Design

Seven (right)

Concept one, the first iteration’s baseline design, presented a number of advantages that

caused it to be worthy of further investigation. Concept one is configured similarly to a

conventional business jet, thus providing a pre-optimized platform for technological

improvement and analysis. Through the evolution of aircraft design, this conventional

configuration has been refined to the extent that a reinvention of the wheel may not be required

to meet the desired end goals. In other words, the application of new technologies and additional

improvements to current aircraft configurations may be able to meet the required design goals

with minimal developmental costs and a conservative amount of risk. Concept one requires no

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additional control surfaces, and the aft fuselage engine mounting allows for a wide variety of

engine choices. In addition, this mounting location would facilitate an unducted fan by keeping

internal cabin noise and vibration to a minimum. Concept one’s T-tail configuration, which can

be seen in Figure 4, can provide a necessary amount of horizontal stabilizer and elevator surface

area while still being able to facilitate an aft mounted unducted fan.

Additional benefits of concept one include a low wing design that allows for minimal

spar interference with the main cabin, shorter and lower weight landing gear struts, ease of

maintenance, and can serve as an emergency exit platform. Concept one’s circular fuselage also

facilitates cabin pressurization and eliminates the need for any special structural reinforcements.

A circular cross section also allows for a large interior cabin volume and adequate storage area

with no increased development costs. Concept one also provides adequate measures of stability,

and incorporates the wide static margin common to current business jets. A large number of

windows can be incorporated into this concept, and its appearance is both attractive and familiar

to the end user.

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Figure 4: Concept One Three View

Concept two also presented a number of advantages warranting further analysis. Sharing

some conventional aircraft characteristics, but capitalizing on many innovative configuration

changes, concept two promises potential improvements effectual of a more pioneering design

toward the system’s overall goals. Similar to concept one, concept two may incorporate several

new technologies to meet the desired mission goals, but also requires higher development costs

and is associated with a higher amount of risk. As in concept one, concept two features an aft

fuselage engine placement that enables a wide variety of engine choices while reducing internal

cabin noise. Though requiring additional structural weight, concept two’s design specifically

addresses environmental noise through the placement of wings below the entirety of the engines.

Lateral engine noise is also shielded using twin outboard-located vertical stabilizers. Again as in

concept one, concept two shares the same advantages of a low wing configuration with a circular

fuselage. The same material technology factors that can be utilized on concept one are equally

applicable to concept two. Unique to concept two is the potential cruise drag reduction

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associated with a lifting canard, which can benefit the environment through reduced fuel

consumption.

Concept two’s development costs promise to be higher than concept one, but are again

variable based upon the degree of improvement desired and number of new technologies

incorporated. The use of a lifting canard requires a more aggressive design process with regard to

stability, and may potentially contain a smaller static margin. Window placement in concept two

is similar to concept one, however the cabin view through the windows in concept two provides

a clear view of the ground from nearly the entire cabin. Concept two’s swept wings, lifting

canards, and twin vertical stabilizers provide it with a futuristic yet familiar set of lines that are

visually stimulating to end users, as shown by the conceptual sketches in Figure 5.

Figure 5: Concept Two Three View

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ADVANCED TECHNOLOGIES

The two advanced technologies that were investigated were an unducted propfan for a

propulsion system and use of composite materials for the structure of the aircraft. The reduction

in fuel consumption is just one of the many goals proposed by NASA’s subsonic fixed wing

program. NASA has set four fundamental goals referred to as N+2 for a subsonic fixed wing

business aircraft set for production in the 2020 timeframe. The four goals are to achieve a noise

reduction down to 42 dB below stage 4 certification, 75% reduction in NOx emissions, 40%

reduction in fuel consumption, and a performance field length reduction of 50%.1 The first three

design goals will hopefully be met by using an innovative propulsion system. An unducted

propfan is an advanced engine design that would incorporate two counter rotating fans that

would be directly connected to the engine’s turbines. General Electric explored the unducted

propfan concept in the 1980s and 1990s and even flew a design named the GE36 on a Boeing

727 test aircraft.2 The project had problems with noise levels and vibration due to the wave drag

created by the high speed fans. The noise levels and plummeting fuel costs of the 1990s caused

the cancellation of the project. Currently, Rolls-Royce, General Electric, and NASA are working

together to achieve a noise reduction level in the unducted propfan concept. Robert Nuttal, Rolls-

Royce’s vice president of future programs strategic marketing, says that “The outcome of this

work is that we are now confident that open-rotor-powered aircraft will be quieter than any

equivalent aircraft flying today and that it will comfortably meet Stage 4 noise legislation.”2 The

unducted propfan technology is the most promising to meet NASA’s N+2 goals within the

specified time frame. In fact, NASA’s environmentally responsible aviation program (ERA) is

devoting much of its research in the subsonic fixed wing project to the unducted propfan

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technology. Guy Norris states “ERA is focused on the goals of NASA’s N+2, a notional aircraft

with technology primed for development in the 2020 time frame as part of the agency’s

subosonic fixed-wing program.”2 Nuttal was quoted in Aviation Week’s December 14, 2009

issue as saying, “So far the GE-NASA experience seems to echo that of Rolls-Royce. We are

able to confirm that the fuel burn will be 25-30% better than today’s products. And, because of

the engine cycle of the open rotor; the nitrous oxide will be 20% lower than another engine with

an equivalent combustor technology. We are now preparing for the next tranche with the next

build of the rig taking place in Q2 2010”.2 Because of the current development being made on

the concept and NASA’s faith in the technology, the team feels it is an appropriate decision to

anticipate using unducted propfans as a propulsion system for the design project.

On the other hand, there is current engine technology from Pratt & Whitney in the form

of a geared turbofan that shows promise. More importantly, the technology is slated for

production before the 2020 production date for N+2. 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- 24,000 lbf depending on engine selection. The thrust range is ideal for our aircraft

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

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The other possible advanced technology will be 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 designs

that are currently in production. The technology is likely to keep improving and deliver 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.

SIZING CODE

Constraint Diagrams

Two separate constraint diagrams, one for each design concept, were used in this portion

of the design process. The constraint diagram associated with the conventional configuration

changed very little since the System Requirements Review; the only significant changes were

updating plane geometry inputs such as aspect ratio and adjusting landing and takeoff distances

through design iterations. A second constraint diagram was developed for the canard pusher

configuration. All the necessary plane sizing inputs were changed, but many values remained the

same. For example, values for CLmax at landing and takeoff are the same for both concepts

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because we have assumed that these values are limited primarily by the flaps included on the

plane; both configurations currently use single slotted flaps. Future analysis will consider adding

Fowler flaps to determine if they are beneficial. Our design group currently does not definitively

know how adding a canard will change many of the values associated with the canard pusher

configuration. Because of that, many values have been assumed to be the same as the

conventional configuration until further differentiation can be achieved. Table 3 lists some inputs

for both configurations.

Table 3: Aircraft Input Values

W0/S TSL/W0 (CL)maxT/O (CL)maxLanding CD0 e AR

Conventional 76 .33 1.5 2.0 .018 .8 8

Canard-Pusher 84 .33 1.5 2.0 .018 .8 10.5

Some values, such as CD0 and SFC, are calculated many times in the sizing code at

different flight velocities and thrust settings. Because of that, theses values vary widely

depending on the location of interest in the flight mission. The corresponding values reported in

the above table are representative of the appropriate flight condition.

It is also clear from this table that both configurations have identical input values for

some significant variables. This is based on an early design-phase assumption necessitated by a

lack of knowledge that both configurations will have similar constraining variables. It is

anticipated that many more differences will exist between the two models as progress is made in

how to mathematically and systematically represent the differences between the two

configurations. Also the engine type will be differentiated between the geared turbofan and the

unducted turbofan jet engines. Both have varying benefits with regards to fuel efficiency, noise,

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and emissions. Future additions to the sizing code will attempt to quantify these variations of the

two engines to see which is most beneficial overall to both aircraft concepts.

Figure 6 below shows the current constraint diagram for the conventional configuration.

This model is limited by second segment climb and landing ground roll. Figure 7 shows the

current constraint diagram for the canard pusher configuration, which is limited primarily by

landing ground roll, takeoff ground roll, and second segment climb. The first two columns of

Table 3 above summarize the initial starting points used in the code for the thrust to weight ratio

and wing loading. The wing loading for both configurations has changed since the system

requirements review, in which the wing loading was estimated at 100 lb/ft2. Both values have

dropped based upon research of exiting planes and because of iteration results in the design

process.

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Constraint Diagram of Conventional

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

40 60 80 100 120 140

W0/S [lb/ft2]

T SL/W

0 top of climb (1g steady, level flight, M= 0.85 @ h=42K, service ceiling)subsonic 2.5g manuever, 250kts @ h=10Ktakeoff ground roll 4700 ft @ h = 5K,+15° hot daylanding ground roll 3500 ft @ h = 5K,+15° hot daysecond segment climb gradientabove h = 5K, +15° hot day

Figure 6: Constraint Diagram for Conventional Configuration

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Constraint Diagram of Canard Pusher

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

40 60 80 100 120 140

W0/S [lb/ft2]

T SL/W

0 top of climb (1g steady, level flight, M= 0.85 @ h=42K, service ceiling)subsonic 2.5g manuever, 250kts @ h=10Ktakeoff ground roll 4700 ft @ h = 5K,+15° hot daylanding ground roll 2950 ft @ h = 5K,+15° hot daysecond segment climb gradientabove h = 5K, +15° hot day

Figure 7: Constraint Diagram for Canard Pusher Configuration

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Sizing Methods

A detailed Matlab script has been developed in order to size both configurations. This

code is used to estimate the size of the aircraft required to complete our design mission, which is

provided in Figure 1. The code takes in approximately one hundred inputs for each

configuration, utilizes a drag calculation at each step of the mission, predicts the empty weight

with a component build up method, and incorporates an early but detailed engine model. While

the code is functioning and provides reasonable estimates, it is under constant revision in an

attempt to improve estimations, to make assumptions more accurate, and to further differentiate

between the two configurations. The general process of the sizing code is shown below in

Figure 8, which displays the solution process graphically in a flow chart.

Figure 8: Flow Chart for Sizing Code

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After every significant change, the sizing code is calibrated twice, once for each

configuration. The conventional model is calibrated against the Gulfstream G550 and the canard

pusher model is calibrated against the Beechcraft Starship. Both of the chosen calibration

models were used because their configurations closely match the design configurations our group

has chosen. Calibration is carried out by running our sizing code with all inputs describing the

corresponding calibration aircraft. The code’s predictions of empty weight, fuel weight, and

gross weight are then compared to published data for each calibration aircraft, and the

appropriate calibration factors are calculated. The calibration factors for the G550 are used in

sizing our conventional model and the calibration factors for the Starship are used in sizing our

canard pusher model. The final calibration factors are listed below in Table 4.

Table 4: Final Calibration Factor Values

Canard Conventional

Empty Weight 0.77 0.77

Fuel Weight 1.33 1.28

Gross weight 1.05 1.01

The calibration factors shown above are larger than desired. This is most likely caused

by two significant sources of error. The first issue in the current sizing code is that the

component-build-up equations used best describe cargo and transport planes from the 1970’s.

This will likely cause some degree of error, in that a business jet in 2020 will have many

significant differences. Additionally, the engine model in use attempts to scale a 50,000 lbs

thrust engine down to approximately 15,000 lbs of thrust. This process, described later in the

report, will also likely cause some degree of error. As the code is continually refined and made

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more and more accurate, it is hoped that the calibration factors will decrease in magnitude and

become continually closer to one.

It is also important to note that both calibrations yield the same value for the fuel weight.

This is because the Starship, which is used to calibrate the canard pusher configuration, uses a

turboprop engine while our code predicts the engine performance of a turbofan or unducted fan.

Because the turboprop engine is so fundamentally different from what the codes analyzes, the

G550 calibration is carried out first, and it is assumed that the fuel weight calibration factor for

the Starship will be the same as the G550. So after the Starship is calibrated, its fuel weight

calibration factor is replaced with the fuel weight calibration factor of the G550.

The inputs for each of our two configurations are provided in Appendix A and Appendix

B. The variables have been sufficiently commented such that the variables and their values

should be self explanatory. A final weights table for all components of our two planes has been

provided below as Table 5.

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Table 5: Final Weight Values

  Canard (lbs) Conventional (lbs)   Canard (lbs) Conventional (lbs)

Wing 5,367 6,232 Flight Instruments 260 254

Canard/Horizontal Tail 728 753 Hydraulics 179 171

Vertical Tail 879 662 Electrical System 943 1,125

Fuselage 7,445 7,392 Avionics 471 394

Main Gear 922 986 Installed Avionics 359 359

Nose Gear 107 162 Air Conditioning 171 466

Nacelle 1,628 1,656 Anti-Ice 147 145

Engine Controls 90 98 Crew 800 800

Engine Starter 130 130 Payload 1,760 1,760

Fuel System 352 360 Fuel Weight 27,300 28,383

Flight Controls 683 766 Empty Weight 39,742 41,257

Installed APU 440 275 Takeoff Gross Weight 72,944 71,365

Stability and Control

It is important for any business jet to be statically stable. Even at the early design stage,

our sizing code predicts the static stability of each configuration very simply by calculating the

static margin. The static margin measures the distance between the aerodynamic center and the

center of gravity of the airplane and normalizes this distance relative to the chord of the wing. In

order to be statically stable, the static margin must be positive. The sizing code calculates the

static margin according to the equation listed below.

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SM=X AC−XCG

Cmac

This calculation is carried out as one of the final steps in the code because the final sizes

and locations of each component must be known to accurately find the center of gravity. If the

location of any component cannot be calculated, it is given directly as an input. The

aerodynamic center for the conventional model is assumed to be at the quarter chord. The

aerodynamic center for the canard pusher is slightly more complicated since it is not explicitly

known how much the canard will affect its location. The effect of the canard on the aerodynamic

center will also depend greatly on whether or not the canard is a lifting canard. For now, we

have assumed that our canard pusher configuration includes a lifting canard, and that this will

cause the aerodynamic center to shift from the quarter chord point of the main wing to the

leading edge of the main wing. With these assumptions, the static margin and center of gravity

have been calculated for each configuration. These values are listed below in Table 6. It is

important to note that the center gravity is listed as a percentage of the fuselage length while the

static margin is listed as a percentage of the chord of the wing. As expected, the center of

gravity for the canard pusher model is much further aft. Because of this, it also makes sense that

the static margin is smaller for the canard pusher model.

Table 6: Static Margin and Center of Gravity

Center of Gravity (% Fuselage Length) Static Margin (% Cmac)

Conventional 51% 37%

Canard 74% 29%

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Additionally, the code also calculates the general sizes of the wing and tail surfaces.

These calculations are based primarily on the aspect ratio, the wing loading, and the tail volume

coefficients. The tail volume coefficients are provided in Daniel Raymer’s textbook and are

based primarily on historical data. It was assumed that the canard can be sized using the

horizontal tail volume coefficients from Raymer’s textbook.4 The current sizes for the wing,

horizontal tail, vertical tail, and canard surfaces are listed below as Table 7.

Table 7: Wing Sizes

Wing Vertical Tail Horizontal tail Canard

Canard

AR 10.5 2.4 - 9

b 96.1 28.1 - 40.7

c 10.5 10.5 - 4.5

S 876.5 329.5 - 183.6

Conventional

AR 8.0 1.0 2.0 -

b 87.2 13.7 23.8 -

c 12.0 12.3 8.9 -

S 950.7 186.6 284.2 -

Drag Component Buildup

The drag for the aircraft may initially be estimated by considering component

contributions. Essentially, the aircraft may be broken down into a handful of major sections with

each section’s drag estimated through a combination of theoretical and empirical equations.

Each component’s drag may also be modeled as a sum of different types of drag.

25

There are four major types of drag being considered in our analysis. The first is parasite

drag, Dp. Parasite drag is comprised of skin friction drag, pressure drag and interference drag.

This drag is a result of surface interactions with a viscous fluid, in this case, air. The coefficient

of parasite drag is defined as:

CD p= ∑

i=1

¿ comp K iQi C f iSwet i

Sref+k CL

2

However, the second term will be used in a different manner described later. Sref is the

reference area, in this case the wing planform area. K is the form factor, a function of

component shape that takes into account pressure drag. Q is interference factor. Depending on

the placement of the component, Q may range from 1 to 1.5. In this case of this aircraft, Q will

be usually be 1. Swet is the wetted surface of the component, described as being the surface area

of the object which would be wetted by dropping the aircraft into a bucket of water.

The second major type of drag is induced drag, D i. Induced drag is a product of finite-

span airfoils, whereby wingtip vortices flow from the lower to upper surface of the wing,

lowering the airfoil’s effective angle of attack. The presence of wingtip vortices leads to a

component of the lift acting as drag since drag is defined as the force produced parallel to the

free stream. The induced drag coefficient is defined according to the equation below.

CDi=

C L2

πAR µ

Here, µ is the span efficiency factor, defined to be 1 when the wing produces an elliptic lift

distribution. Wing shape is very important in determining the lift distribution and many aircraft

employ near-elliptic lift distributions. Current general aviation aircraft generally have an

efficiency factor ranging between 0.75 and 0.85. The aircraft’s induced drag coefficient is

currently being estimated using an efficiency factor of 0.80.

26

The third type of drag is wave drag, Dw, also known as compressibility drag. Wave drag

results from the formation of shock waves produced around the aircraft at sonic or near-sonic

speeds. The energy of the shock waves radiating off the aircraft is perceived as drag. While the

aircraft will not ever fly at Mach 1, localized airflow will most certainly reach sonic speed. As

of yet, no method for predicting or estimating wave drag has been found or included for this

application, thus it is currently assumed that the aircraft experiences 20 counts of wave drag.

While likely incorrect, the inclusion of a constant value of wave drag serves as a place holder for

future development and as a rough estimate for the early design stage analysis. This omission

will be corrected in the future.

The last type of drag being considered is miscellaneous drag, Dmisc, which is a factor

taking into account leakage, protuberance, and imperfections in manufacturing. This drag

accounts for issues such as gaps in control surfaces, external lighting, antennae, rivets and other

surface blemishes. Based off of sample drag prediction information, conservative estimates put

the aircraft’s miscellaneous drag coefficient at 5% of the aircraft’s parasite drag coefficient. This

approximation is used exactly in the current drag prediction.

To vary slightly from the described list of drag coefficients, it is convenient to redefine

two aspects of the total drag into a new term called ‘drag due to lift’. This term is comprised of

the parasite drag term dependent on CL, kCL2, and the induced drag term, CL

2

πAR µ. Together these

terms are combined to produce the term CL2

πARe, the drag due to lift, where e is Oswald’s

efficiency factor. As stated previously, e is estimated to be 0.8. For the purpose of this report it

is reasonable to determine CL at cruise based off of the assumption that lift equals weight during

cruise. Accordingly, Wcruise = CLqS. Sample calculations are provided in APPENDIX C.

27

Now that the types of drag have been discussed, it is time to consider the contributions

from the aircraft components themselves. The major components being considered on the

aircraft are the fuselage, main wing, vertical and horizontal tails, canards, nacelles and pylons.

The fuselage is currently being modeled as a cylinder with a nose and tail section added on. The

fineness ratio of the aircraft, λf, was determined via a combination of current similar aircraft’s

fineness ratios and mission design criteria such as ceiling height and volume per passenger.

The wetted area of the fuselage is a function of the fineness ratio, and can be calculated

according to the following equation.

Swet f=π Df l f (1− 2

λ f )23 (1+ 1

λ f2 )

Here, Df and lf are the diameter and length of the fuselage, respectively. The wetted area is

necessary to calculate the parasite drag, as skin friction only affects the wetted area of the

aircraft. The wetted area for the nacelles uses a simple model of a cylinder. For this the wetted

area is computed with the following equation.

Swet N=π DN lN

Similarly to the fuselage D is the diameter and l is the length. Since the wings, horizontal tail,

vertical tail, and pylons all all similar shapes, it is possible to use the same equation. For these

surfaces, the wetted area was calculated with the following equation.

Swet=Sw 2c

Sw is the plan form area, c is the curvature of the surface, and multiplied by 2 to take the top and

bottom surface into account.

For the component drag build up, the form factor for the shape of each component was

needed. The form factor is shape dependent and varied for each component according to the

equations listed below.

28

K f =1+ 60λ f

3 +λ f

400

K N=1+ 0.35

(lN

DN)

Kw=1+Z ( tc )+100( t

c)

4

In these equations, the subscript f is for the fuselage, N is for the nacelle, and W is for the wings.

The equation for the wings was also used for the pylons, vertical tail, and horizontal tail. Here, λf

is again the fineness ratio, Z is the sweep correction factor, and t/c is the thickness to cord ratio.

To find this form factor, the sweep correction factor is needed. This factor was found using the

following equation.

Z=(2−M 2)cos Λc /4

√1−M 2(cos Λc /4)2

With the form factor and wetted area known, the last required piece is the interference

factor. The interference factor is determined by where different components are located relative

to each other. For this aircraft the nacelles have an interference factor with the fuselage. This

factor is based on the distance from the fuselage to the nacelle. If the distance is greater than the

diameter of the nacelles, the interference factor, Q, equals one. If the nacelle is resting directly on

the fuselage then the interference factor, Q, is 1.5. The wing and the fuselage also interfere with

each other. The position of the wing, being a low wing with fillets, results in an interference

factor that is also 1. For the pylons and tail, it is similar to a midwing set up, which will result in

an interference factor also of 1. With these four factors all known, it is possible to use the

following equation to sum up the different components and their drags to form a total parasite

drag.

29

CD p= ∑

i=1

¿ comp K iQi C f iSwet i

Sref

The total drag coefficient, CD is very simply defined to be a sum of the four types of drag

discussed previously.

CD=CD p+CDi

+CDmisc+CD w

CD=¿ ∑i=1

¿comp K i QiC f iSwet i

Sref + CL

2

πARe + CD misc + CD w

ENGINE SELECTION

Two separate and distinct engine concepts are being explored by the team: an unducted

propfan and a geared turbofan. Both concepts offer significant savings in both emissions and fuel

consumption, and incorporate the most modern designs in the aircraft propulsions industry. The

details of each particular engine model were discussed previously in the Advanced Technologies

section of this report, but a table of the estimated performance improvements is provided below

as Table 8.

Table 8: Engine Model Projection and Comparison

Predicted Engine Performance by 2020

 % Fuel

ReductionNoise Reduction

(dB)% Emission Reduction

Unducted fan 25-30 Meets Stage 4 20

30

Geared TurboFan 15 20 50

An in depth analysis process is currently being used to model the engine as part of the

sizing code. The main purpose of the engine model is to predict how much fuel will be required

to complete the design mission, and more specifically the fuel weight. This process starts with

the engine data provided in Raymer’s textbook for a “High Bypass Turbofan.”4 Unfortunately,

Raymer is not more specific than that as to what model the engine actually is. The engine data

plots are scanned directly into a MATLAB script which converts the data to tabular form and

interpolates where necessary. This engine is much larger than required, and must be scaled

down from 50,000 lbs of thrust to approximately 15,000 lbs of thrust in order to be useful. The

scaling process is performed simply by calculating a scale factor, which equals the desired

amount of thrust at any given time divided by 50,000 lbs, and multiplying the thrust data by this

value. As engines shrink in size and performance, however, their fuel consumption and

efficiency generally worsen. Because of this, the specific fuel consumption was scaled inversely

proportional to the established scale factor.

With the data provided and the scaling operations performed, the fuel consumption can

then be predicted as a function of Mach number, altitude, and throttle setting. Different

calculations are used for each segment of the flight mission: taxi, takeoff, climb, cruise, descent,

loiter, and land. The appropriate equations provided by Professor Crossley in lecture were used

exactly for each mission segment. Some sections are simple and straight forward, such as taxi,

takeoff, and landing, while the others require some looping to find optimum climb rates or lift to

drag ratios. The climb calculations predict the fuel consumed at a velocity of 250 knots up to

31

10,000 ft followed by the fuel consumed at best rate of climb up to the initial cruise altitude of

41,000 ft. The calculations are broken down into 500 vertical foot increments in which it is

assumed that atmospheric properties are constant. The cruise section is broken into twenty

equal-portion increments, and it is assumed that cruise altitude will remain constant at 41,000 ft.

The aircraft is now also assumed to be flying at its maximum lift to drag ratio. This same

performance assumption is used for loiter calculations. Descent is given a zero-range credit,

which has not changed since the System Requirements Review.

Each engine is analyzed using the above process with the technology factors from Table

8 included. When research states that a particular engine will decrease fuel consumption by X%,

that factor is used directly when applied to the specific fuel consumption value in calculations.

For example, the unducted fan reportedly reduces fuel consumption by 30%, so all SFC values

taken from the engine model in Raymer’s textbook are scaled and then multiplied by 0.70.

While noise emission is not currently predicted, this same method will be used when the

calculations become possible.

32

SUMMARY OF AIRCRAFT CONCEPTSThe major goals of this design are to provide an environmentally conscious, ultra-long

range corporate aircraft, which meets NASA’s N+2 goals, is fast enough to be marketable, and

can operate out of general aviation airports. Over the course of the design process, these broad

goals have given way to a detailed set of requirements. These requirements are shown in the left

column of the requirements compliance matrix in Figure 9. In addition to specifying the name of

the criteria being measured, this matrix also specifies the target value of each criterion, the

minimum threshold value necessary to satisfy each requirement, and the current value based on

the present state of the design. It is important to note that several of the current values shown are

estimates, and may not be constrained by the current level of system detail. These estimated

values are denoted by a trailing asterisk in the matrix.

The range, cruise mach, and initial cruise altitude requirements were derived from the

design mission. While the locations of the design mission have not changed, the still-air range of

the mission has been adjusted since the System Requirements Review (SRR). Originally, the

aircraft’s design mission range was specified as the great circle distance between Los Angeles

(LAX) and Hong Kong (HKG). This range ignored the fact that strong headwinds would likely

be unavoidable when traveling west to Hong Kong over the Pacific. As a conservative figure, a

headwind of 60 knots was assumed for the duration of the flight. This resulted in an equivalent

still-air range of 7100 nautical miles. A threshold value of 6960 nautical miles was then

determined from a less conservative, but still realistic, headwind of 50 knots. The cruise mach

and initial cruise altitude remain on target since the SRR, and are incorporated into the aircraft’s

current sizing and weight estimates. The target and threshold values of these requirements are

both based on producing an aircraft that is fast enough to be marketable. The cruise mach is

33

representative of most modern jet aircraft; and the initial cruise altitude is high enough to allow

for flight above most of the cruise altitudes used by part 121 carriers.

The maximum passenger capacity and the volume per passenger per hour requirements

were obtained from the aircraft’s cabin layout and design mission. The maximum number of

passengers has remained on target since the SRR, and is factored into the cabin’s seating

configuration. The threshold value of 8 listed in Figure 9 signifies the number of passengers

carried on the aircraft’s design mission. The volume per passenger per hour requirement was a

new system requirement added to provide a measure of passenger comfort. The current threshold

value of 2.28 ft3/(pax·hr) represents the minimum level necessary to ensure the design mission is

not arduous to passengers.5 The current and target value of volume per passenger hour is 13.3

ft3/(pax·hr). This value was calculated from a CATIA model of the aircraft’s cabin, and is

considered to be comfortable over the course of the design mission, which is approximately 13

hours.5

The balanced takeoff field length at maximum takeoff weight (MTOW) is also a new

requirement, which replaced the previous and more ambiguous requirement of “takeoff distance”

featured in the SRR. Balanced takeoff field length is the minimum length of runway required

such that a twin engine aircraft can operate safely with one engine inoperative during any stage

of takeoff. It was chosen as the metric for the design’s takeoff performance, since it specifies the

minimum amount of runway necessary to comply with safety regulations. The target value of

6000 ft was selected since it is close to the length of runways found at many general aviation and

regional airports. The threshold of 7000 ft was chosen since it was approximated as the

maximum required field length at which the aircraft could still be marketed as having reasonable

airport access. Since the current level of design detail was not sufficient to calculate balanced

34

field length, the ground roll values for each concept were included in the matrix. These ground

roll values were calculated using the constraint diagrams for each concept. While the current

ground roll values are promising in comparison to existing aircraft, this requirement will

continue to be of great concern until field length can be approximated.

The cabin noise, cumulative noise, emissions, sill height, and variable cost requirements

remain unchanged since the SRR, and are currently only estimates unconstrained by the present

level of design detail. Reasonable cabin noise values were estimated from existing aircraft data.

However, it is not yet possible to calculate cabin noise, or the difference in cabin noise that is

expected to exist between the two concepts. Cumulative noise limits were also estimated from

existing aircraft data, while emissions levels were placed relative to the NASA N+2 goals.

Cumulative noise limits cannot yet be calculated, but emissions levels will be able to be refined

once more detail on the power plants is obtained. Sill height was estimated based on the need to

facilitate aircraft loading, and also by studying the sill heights of existing business jets. This is

expected to be governed by takeoff rotation considerations in the future. Variable costs were

estimated from existing aircraft cost data, and the cost of operating a current aircraft with a

similar design mission was used as the threshold. Clearly, the variable cost cannot be calculated

at this point in the design.

Lastly, the cruise specific range requirement represents the desired fuel efficiency of the

design. The threshold value of 0.26 nm/lb was based on existing aircraft performance.6 Current

values were calculated for each concept, using the fuel weight determined from the sizing code

and the range of the design mission. This method was a crude approximation, but a more

accurate answer will be achievable once more power plant details are determined. The target

35

value was set to the range necessary to fly the design mission with reserves, assuming current

fuel weight.

Figure 9: Requirements Compliance Matrix.

Scaled models of each concept have been developed in CATIA. Three view drawings

with a few general sizing dimensions and a three dimensional screen shot from CATIA have

been provided for the Canard Pusher Configuration as Figure 10 and Figure 11. Similar plots

follow for the Conventional Configuration in Figure 12 and Figure 13.

36

Figure 10: Three View Drawing - Canard Pusher Configuration

Figure 11: Isometric CATIA Screen Shot - Canard Pusher Configuration

37

Figure 12: Three View Drawing - Conventional Configuration

Figure 13: Isometric CATIA Screen Shot - Conventional Configuration

38

SUMMARY

The second phase of this project consisted of developing many designs using various

technologies and configurations and then to narrow those designs down to the two best choices

for further analysis. The two choices that best met the developed requirements are a conventional

aircraft and an aircraft with canards and an aft swept back wing. Using these two configurations,

advanced technologies will be implemented such as composites and more efficient engines like

the unducted turbofan or the geared turbofan. Current calculations show that the aircraft, sized to

meet the demands of the target customer and designed to achieve the environmental goals, are in

the low 70,000 lb range. Further analysis will be done to evaluate these two designs to determine

which is best and which meets those requirements best.

NEXT STEPS

Now that Team 2 has finished the System Design, there are many more steps that need to

be taken before the conceptual design of the aircraft can be completed. For the wing, tail, and

canard, airfoils need to be selected to better predict the drag polar and general performance of

each piece. In propulsion, the engine selection between the geared turbofan and unducted fan

needs to be decided. In structures, a lot of work needs to be done especially in loading of

different parts and in integrating the landing gear. The longitudinal stability of each aircraft

needs to be finalized and the lateral stability needs to be analyzed for cross-wind landing and one

engine out conditions. Using all of this updated data, the best performance data needs to be

calculated including all airplane weights and sizes. The internal and external noise of each

aircraft needs to be assessed. Finally the estimated cost of each aircraft need to be found. Using

39

all of this and our component sizing code, carpet plots need to be used to graphically optimize

each of our designs. Once all of the calculations and optimizations have been completed a final

aircraft selection will be made.

40

REFERENCES1“Subsonic Fixed Wing Project”. NASA. 08 February 2010.

http://www.aeronautics.nasa.gov/fap/sfw_project.html

2 Norris, Guy. “Rotor Revival”. Aviation Week & Space Technology. 14 December 2009. pages

54-57.

3 "PurePower PW1000G." Home. Web. 30 Mar. 2010.

<http://www.pw.utc.com/Products/Commercial/PurePower+PW1000G>.

4 Raymer, Daniel P. Aircraft Design A Conceptual Approach (Aiaa Education Series). New

York: AIAA American Institute of Aeronautics & Ast, 2006. Print.

5Torenbeek, Egbert. Synthesis of subsonic airplane design an introduction to the preliminary

design, of subsonic general aviation and transport aircraft, with emphasis on layout,

aerodynamic design, propulsion, and performance. Delft: Delft UP, Nijhoff, Sold and

distributed in the U.S. and Canada by Kluwer Boston, 1982. Print.

6Jane's All The World's Aircraft. Web. 11 Feb. 2010.

<http://jawa.janes.com/public/jawa/index.shtml>.

41

APPENDIX AConventional Configuration Input Variables

42

%% Input section % aircraft contant ratiosWo_S = 76; %lbs/ft^2 from initial contrait diagram Cfactor = GetCfactors('G650'); %Tech Factors engine.tech_fact = .7; %Technology factor - SFC (~30%) (~40%=projected) % wing variableswing.AR = 8;wing.taper = .3;wing.sweep = deg2rad(35); %radwing.t_c = .08;wing.Scs = 25; %control surface area on wing (ft^2) %vertical tail variablestail_vt.Cvt = .09;tail_vt.Lvt = 40; %feettail_vt.taper = .8; tail_vt.sweep = deg2rad(25); %radtail_vt.t_c = .08;tail_vt.Ttail = 1; %1 for t-tail, 0 otherwisetail_vt.AR = 1;tail_vt.Scs = 40; %control surface area on vertical tail (ft^2) % canard variablescanard.Lc = 40; %feetcanard.taper = .5;canard.Cc = 1;canard.AR = 2;canard.sweep = deg2rad(40); %radcanard.t_c = .08;canard.Kuht = 1; %all moving = 1.143, otherwise 1canard.Fw = 9; %fuselage width at canard location (ft)canard.Scs = 50; %control surface area on canard (ft^2) %pylon/Nacellesnacelle.L = 20;nacelle.D = 5;pylon.c = 1;pylon.t_c = .3;pylon.S = 10; %fuselagefuselage.L = 90; %length of fuselage (ft)fuselage.D = 9; %diameter of fuselage (ft)fuselage.SDepth = 7; %fuselage structural depth (diam at wing box) (ft)fuselage.Lc = 40; %wing mac to canard mac (ft) %landing geargear.location = 1; %1 = wing, 1.12 for fuselage mounted

43

gear.Nl = 2.5; %ultimate landing load factor (1.5*Ngear)gear.length_main= 5; %ftgear.length_nose = 5; %ftgear.main_wheels = 4; %number of main gear wheelsgear.main_struts = 2; %number of main gear shock strutsgear.nose_wheels = 2; %number of nose gear wheels % mission variablesmission.num_pax = 16;mission.weight_pax = 220;mission.num_crew = 4;mission.weight_crew = 200;mission.range = 7000; %nmimission.q = 200; %dynamic pressure at cruisemission.Nz = 2.5*1.5; %ultimate load factor (1.5*load factor)mission.Mcruise = .85;mission.cruise_alt = 42000; %engine variablesengine.Kng = 1.017; %1.017 for pylon mounted nacelle, 1.0 otherwiseengine.nacelle_length = 20; %ftengine.nacelle_width = 5; %ftengine.weight = 3000; %lbsengine.number = 2; %total number of enginesengine.d2c = 110; %ft - total distance from engines to cockpit (sum for both engines)engine.Kr = 1; %1.133 if reciprocating engine, 1.0 otherwiseengine.Ktp = 1.0; %0.793 if turboprop, 1.0 otherwiseengine.SFC_cruise = .5; %1/hrengine.SFC_loiter = .4; %1/hr %controls variablescontrols.num_fnx = 5; %number of functions performed by controlls (usually 4-7)controls.num_mech = 1; %number of mechanical functions (usually 0-2)controls.Iy = 1001000; %yawing moment of inertia (lb ft^2) % aerodynamicaero.V_stall = 210; %ft/s - stall for the main wing %weightweight.Wo_guess = 100000;weight.APU_u = 125; %lbs - uninstalled weightweight.avionics_u = 250; %lbs - uninstalled avionics weight %miscallaneousmisc.Rkva = 50; %system electrical rating (kV) - usually 40 to 60 for transportsmisc.La = 250; %electrical routing distance, generators to avionics to cockpit (ft)misc.Ngen = 2; %number of generators (usually equal to number of engines)misc.Vp = 4000; %volume of pressurized section (ft^2)

44

%Locations (note, all in feet relative to nose, plane length ~ 90 ft)loc.apu = 82;loc.avionics = 5;loc.fuel = 42.5;loc.wing = 42.5;loc.canard = 90;loc.tail_vt = 86;loc.fuselage = 45;loc.main_gear = 50;loc.nose_gear = 10;loc.nacelle = 72;loc.engine_controls = 5;loc.engine_starter = 5;loc.fuel_system = 42.5;loc.flight_controls = 5;loc.instruments = 5;loc.hydraulics = 50;loc.electrical = 45;loc.furnishings = 35;loc.airconditioning = 6;loc.antiice = 35;loc.crew = 10;loc.payload = 35;loc.engines = 72;loc.AC = 50;

45

APPENDIX BCanard Pusher Configuration Input Variables

46

%% Input section % aircraft contant ratiosWo_S = 84; %lbs/ft^2 from initial contrait diagram Cfactor = GetCfactors('Starship'); %Tech Factors engine.tech_fact = .7; %Technology factor - SFC (~30%) (~40%=projected) % wing variableswing.AR = 10.53;wing.taper = .2;wing.sweep = deg2rad(30); %radwing.t_c = .1;wing.Scs = 10; %control surface area on wing (ft^2) %vertical tail variablestail_vt.Cvt = .09;tail_vt.Lvt = 23; %feet wing mac to tail mactail_vt.taper = .8; tail_vt.sweep = deg2rad(30); %radtail_vt.t_c = .08;tail_vt.Ttail = 0; %1 for t-tail, 0 otherwisetail_vt.AR = 2.4;tail_vt.Scs = 20; %control surface area on vertical tail (ft^2) % canard variablescanard.Lc = 50; %feetcanard.taper = 1;canard.Cc = 1; %volume coefficient canard.AR = 9;canard.sweep = deg2rad(35); %radcanard.t_c = .08;canard.Kuht = 1.143; %all moving = 1.143, otherwise 1canard.Fw = 9; %fuselage width at canard location (ft)canard.Scs = 35; %control surface area on canard (ft^2) %pylon/Nacellesnacelle.L = 15;nacelle.D = 5;pylon.c = 1;pylon.t_c = .3;pylon.S = 10; %fuselagefuselage.L = 90; %length of fuselage (ft)fuselage.D = 9; %diameter of fuselage (ft)fuselage.SDepth = 7; %fuselage structural depth (diam at wing box) (ft)fuselage.Lc = 50; %wing mac to canard mac (ft) %landing geargear.location = 1; %1 = wing, 1.12 for fuselage mounted

47

gear.Nl = 2.5; %ultimate landing load factor (1.5*Ngear)gear.length_main= 4; %ftgear.length_nose = 4; %ftgear.main_wheels = 4; %number of main gear wheelsgear.main_struts = 2; %number of main gear shock strutsgear.nose_wheels = 1; %number of nose gear wheels % mission variablesmission.num_pax = 18;mission.weight_pax = 220;mission.num_crew = 4;mission.weight_crew = 200;mission.range = 7000; %nmimission.q = 200; %dynamic pressure at cruisemission.Nz = 1.5*2.5; %ultimate load factor (1.5*load factor)mission.Mcruise = .85;mission.cruise_alt = 42000; %engine variablesengine.Kng = 1; %1.017 for pylon mounted nacelle, 1.0 otherwiseengine.nacelle_length = 20; %ftengine.nacelle_width = 5; %ftengine.weight = 3000; %lbsengine.number = 2; %total number of enginesengine.d2c = 100; %ft - total distance from engines to cockpit (sum for both engines)engine.Kr = 1; %1.133 if reciprocating engine, 1.0 otherwiseengine.Ktp = 1; %0.793 if turboprop, 1.0 otherwiseengine.SFC_cruise = .5; %1/hrengine.SFC_loiter = .4; %1/hr %controls variablescontrols.num_fnx = 5; %number of functions performed by controlls (usually 4-7)controls.num_mech = 1; %number of mechanical functions (usually 0-2)controls.Iy = 1000100; %yawing moment of inertia (lb ft^2) % aerodynamicaero.V_stall = 220; %ft/s - stall for the main wing %weightweight.Wo_guess = 75000;weight.APU_u = 200; %lbs - uninstalled weightweight.avionics_u = 300; %lbs - uninstalled avionics weight %miscallaneousmisc.Rkva = 50; %system electrical rating (kV) - usually 40 to 60 for transportsmisc.La = 150; %electrical routing distance, generators to avionics to cockpit (ft)misc.Ngen = 2; %number of generators (usually equal to number of engines)misc.Vp = 737; %volume of pressurized section (ft^2)

48

%Locations (note, all in feet relative to nose, plane length ~ 90 ft)loc.apu = 10;loc.avionics = 5;loc.fuel = 78;loc.wing = 78;loc.canard = 10;loc.tail_vt = 83;loc.fuselage = 45;loc.main_gear = 80;loc.nose_gear = 10;loc.nacelle = 85;loc.engine_controls = 5;loc.engine_starter = 5;loc.fuel_system = 78;loc.flight_controls = 5;loc.instruments = 5;loc.hydraulics = 40;loc.electrical = 45;loc.furnishings = 35;loc.airconditioning = 6;loc.antiice = 70;loc.crew = 10;loc.payload = 35;loc.engines = 85;loc.AC = 70;

49

APPENDIX CDrag Sample Calculations

50

At a cruise height of 42,000 feet, at a cruise speed of 0.85 Mach, q is found to be

q = 0.5*(0.000532 slugs/ft3)*(0.85*968.076 ft/sec)2

q = 180.11 psf

As well, S for the aircraft is currently estimated at 1035 ft2. Thus,

CLcruise = WqS ≈

90,000 lb180.11 psf ∗1035 ft2 = 0.483

The drag due to lift will vary slightly based on whether the canard or standard aircraft

setup is considered, since the canard aircraft boasts an AR =10 while the standard aircraft

currently has an AR=8. Thus, for the standard aircraft,

CD = 0.01924 + 0.01159 + 0.0010+ 0.0020

= 0.03384

For the canard,

CD = 0.01924 + 0.009274 + 0.0010 + 0.0020

= 0.03151

Once CD is found the total drag force, DTotal, may be found very easily by the equation

DTotal = CD q S

Thus, the current total drag for the standard aircraft is found to be

DTotal = 0.03384*(180.11 psf)*1035 ft2

DTotal = 6308 lbf ≈ 6310 lbf

The current total drag for the canard aircraft is

DTotal = 0.03151*(180.11 psf)*1035 ft2

DTotal = 5837 lbf ≈ 5840 lbf

Note that it is reasonable to approximate the drag to the ten’s place based on the amount of

approximation present in the calculations.

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