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ME5539-ME5309 Design and Analysis of Aircraft and

Aerospace Vehicles

AIRBUS Dual role Aircraft Specification 2014

Camilo Vergara

1010295

Abstract-

The purpose of this report is to showcase a proposed conceptual aircraft which has been designed

through the use of AAA software, display the iterative processes followed to get to the

aforementioned design, and ultimately prove that the aircraft is capable of both mission profiles

required in the specification brief.

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Contents 1. Introduction- ....................................................................................................................................... 4

1.2 Current Aircraft ............................................................................................................................. 4

1.3 House of Quality ........................................................................................................................... 5

1.4 Design flow Chart .......................................................................................................................... 7

2. Mission Profile- ................................................................................................................................... 8

2.1 Definition ...................................................................................................................................... 8

2.2 Performance requirements ........................................................................................................... 8

3. Initial Weight Estimation- ................................................................................................................... 9

3.1 Fuel Fractions ................................................................................................................................ 9

3.2 Mission Segment Fuel Mass ........................................................................................................ 10

3.3 Take-Off Weight Estimation ........................................................................................................ 10

3.4 Useful Load for 28 PAX Mission Performance ............................................................................ 12

4. Aerodynamics- .................................................................................................................................. 13

4.1. Lift .............................................................................................................................................. 13

4.1.1 Wing ..................................................................................................................................... 13

4.1.2 Horizontal Tail ...................................................................................................................... 14

4.1.3 Vertical Tail........................................................................................................................... 14

4.1.4 Flaps and Slats ...................................................................................................................... 14

4.2 Drag- ............................................................................................................................................ 15

4.2.1 Drag Polars ........................................................................................................................... 15

4.2.2 L/D Ratios ............................................................................................................................. 16

5. Performance- .................................................................................................................................... 16

5.1 Components ................................................................................................................................ 16

5.1.1 Stall Speed ............................................................................................................................ 16

5.1.2 Take-off Distance ................................................................................................................. 17

5.1.3 Climb .................................................................................................................................... 17

5.1.4 Maximum Cruise speed ....................................................................................................... 17

5.1.5 Manoeuvres/Turn ................................................................................................................ 17

5.1.6 Landing Distance .................................................................................................................. 17

5.2 Matching Plot .............................................................................................................................. 18

6. Geometry- ......................................................................................................................................... 19

6.1 Wings & Empennage ................................................................................................................... 19

6.2 Fuselage ...................................................................................................................................... 20

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6.3 Landing Gear ............................................................................................................................... 22

6.4 Ventral Fins ................................................................................................................................. 23

7. Propulsion- ........................................................................................................................................ 23

8. Stability and Control- ........................................................................................................................ 24

8.1 Horizontal and Vertical Tail sizing ............................................................................................... 24

8.2 Static Margin & Neutral Point ..................................................................................................... 24

9. V-n Diagram- ..................................................................................................................................... 25

10. Conclusion- ...................................................................................................................................... 26

Appendix A ............................................................................................................................................ 27

Appendix B: Mission Profile .................................................................................................................. 28

Appendix C: Weights Estimation ........................................................................................................... 28

Appendix D: Aerodynamics ................................................................................................................... 29

Appendix E: Drag ................................................................................................................................... 32

Appendix F: Performance ..................................................................................................................... 35

Appendix G: Geometry ......................................................................................................................... 36

Appendix H: Propulsion ........................................................................................................................ 39

Appendix I: Stability and Control .......................................................................................................... 40

Appendix J: V-n Diagram ....................................................................................................................... 40

Bibliography .......................................................................................................................................... 41

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1. Introduction-

Since its conception, the Civil Aviation sector has grown into the huge industry generating

over 32 million jobs globally and shifting over 2.2 billion passengers annually [1]. It is constantly

expanding and improving in efficiency. For example since the first jet airliners, the fuel efficiency of

these aircraft has dropped considerably, due to improvements in engine design, and aerodynamics.

Figure 1.0.1 below shows this visually in a graph.

Figure 1.0.1 displaying increase of fuel efficiency since early Civil Aviation [1]

In the past few decades civil airliners have started to focus on developing families of aircraft

capable of multi role functions, essentially covering different mission profiles with the same

economic base aircraft, with a range of capabilities dependent on passenger configurations and any

slight extension to the fuselage (examples include the B737 range: 737-600/-700/-800/-900ER) [2] .

The project undertaken involved a task to design at a conceptual level an aircraft that could

be used as a dual role platform for operations in and out of London City Airport (LCY). This intra-

European and intercontinental potential would be considered for potential future family growth, of

+/- 20% passenger capacity for the different versions.

1.2 Current Aircraft

As a basis to start on, market research was undertaken to assess the trends of current

generation of aircraft which lie close to the specifications of the design project. The aircraft chosen

were samples around the rough figure of between 80 to 120 PAX aircraft.

The graphs comparing several performance values shown above can be seen in the appendix

on figure 1.2.2 and were implemented to get a general sense the design trends of a small sample of

aircraft at either end of the required design spectrum. This was vital in getting the first guess sizing

of the conceptual vehicle prior to calculations and analysis, as the various pieces of information gave

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an approximate limits for current generation aircraft regarding basic geometric and performance

data.

Table 1.2.1 depicts general properties of a range of aircraft covering a wide margin of capabilities.

1.3 House of Quality

A basic house of quality what vs Hows matrix was implemented to help source out

important design factors which should be kept at a higher importance than others. Figure 1.3.1

below shows the results from the analysis.

Take-off Weight

Power plant(s)

Fuel Weight

Useful Load

Take off Distance/ Landing Distance

Cruise Speed

Stall Speed

Wing Area

Aspect Ratio

CG Location

Capable of operating from London City Airport

1 9 1

Fuel Efficiency

3 9 1 3 3 1 3

Fast travel times

3 9

Passenger Capacity

1 3 9 1 1 1 9

Technical Requirement importance

4 13 4 12 10 12 2 1 4 9

Relative Percentage importance

6% 18% 6% 17% 14% 17% 3% 1% 6% 12%

Figure 1.3.1 House of Quality Customer vs Technical requirements relationship (What vs How matrix)

AUW (Kg) at MTOW

Wingspan (m)

Range (Km)

Max PAX

Max Altitude (m)

Target Ceiling (m)

Fuselage length (m)

CRJ- 700LR 34,930 23.2 3700 79 12500 12,500 32.5

E170 LR 37,200 26 3889 78 11900 12,500 29.9

B737-900 85,130 35.7 5990 215 12500 12,500 42.1

ARJ21-700 40,500 27.28 2200 90 11900 12,500 33.46

B 737-600 66,000 34.3 5,970 110 12,496.80 12,500 34.2

A319-100 64,000 35.8 6,850 124 12,500.00 12,500 33.84

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From the House of Quality investigation, the customer preferences was decided from

aspects in the project brief regarding expressed design requirements. The important technical

factors to keep in mind during the conceptual design phase in descending order of importance are:

Power plant/ propulsion choice (18%) - It is critical to use the appropriate sized engine with

suitable fuel efficiency during operation which varies from full throttle during take-off, a constant

low value for cruise, and all the stages in-between and after through to landing.

Cruise Speed (17%) – The velocity at which the aircraft would spend most of its flight on is

critical as it dictates the total travel time between destinations in a long haul journey.

Useful Load (17%) – The ability to carry the stated 100 Passengers in the economy class set

up, as well as the 28 passengers on a business class configuration. This is required by the customers

of the new aircraft.

Take-off/Landing distance (14%) – It is a requirement to be able to land and take off from

London City Airport (LCY), so that is an ultimate design limit. The aircraft must be able operate from

LCY’s Runway, which is 1199 m in length (3933.7 feet) but allows for a take-off and landing distance

of 1319 m (4327 feet) [3].

CG location (12%) - is important for balancing the aircraft as well as dictating the location of

key geometric features such as wings, aft landing gear, vertical and horizontal tail. It is a key

component of the stick free Static margin calculations in the stability module of AAA to achieve a

stable airworthy aircraft.

Other technical requirements worth noting are:

Take-off Weight (6%) – Gross weight of the maximum load the aircraft can carry. Within

reason it must be adequately small for an appropriate aircraft of its class (100 PAX) to minimize any

detrimental effects the extra weight could add to fuel efficiency.

Aspect Ratio (6%) – This affects the lift generated on the wing which is important to keep it

at relatively high values, using existing aircraft wing configurations as a guideline.

Stall Speed (3%) – This is a design aspect that has a knock on effect on the approach speed

and the distance required to land when all high lift devices are deployed. As per the brief

description, the approach speed is limited to below 130 knots, which is 1.3 times the velocity of stall.

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1.4 Design flow Chart

In an aircraft design project it is important to make sure to understand and follow logical

steps so as not to get lost in-between all the design work, and keep on track to work towards a

finished product which abides by any design requirements set at the beginning of the project. An

example of this process is shown below on figure 1.4.1.When outlined as a visual reference to work

to, it clearly shows how at all points of the first conceptual stage, the design team must be

constantly checking back and making sure that thus far they still adhere to any requirements set out

at the beginning.

Figure 1.4.1 Aircraft Conceptual Design Process [4]

As a result of such methodologies, it was important to use appropriate software which

follows the same steps through the conceptual design phase. From the figure above it is important

to stress that aircraft design is an iterative process, and all the data presented in this report

represents the final iteration of the aircraft at a conceptual level.

AAA 3.5 Software (Advanced Aircraft Analysis) was used in aid of the aircraft design set for

this project as it was built around Roskam’s Airplane Design books which form the basis on which to

fully design an aircraft from concept through to preliminary and finally detailed phases later on. The

program allows for the conceptual and preliminary design of wide selection of different types of

New Concept Ideas Design Requirements

Technology

Availability

Concept Sketch

First-guess sizing

Initial Layout

Initial analysis

Aerodynamics

Weights

Propulsion

Sizing & performance

optimization

Revised Layout

Analysis

Aerodynamics

Propulsion

Stab & control

Structures

Cost

Subsystems

ETC

Refined Sizing &

performance optimization

Preliminary design

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aircraft. As such a copy of AAA files created and used in the process of this aircraft design project will

be attached via a CD to this report for cross referencing purposes.

2. Mission Profile-

2.1 Definition

Figure 2.1.1 Mission Profile for the aircraft as required by the project brief

Above shows a diagram which sums up the stages during the flight profile required. From

this they were split into individual sections and utilized to get the fuel fractions through the weights

estimation module of AAA. This is linked with a table in appendix B organising the mission profile

into sections of time, distance and fuel.

2.2 Performance requirements

The mission required is split into separate phases in accordance with figures 2.1.1 and 2.1.2.

It initiates with the start-up and taxi out phase of the mission, which is expected to last 7 minutes.

From here the next phase is the take-off, initial climb to 1,500 ft and acceleration to 250 knots.

Upon reaching 1500 ft the third phase starts which is the climb to 10,000 ft at 250 knots, at

which point a secondary climb would begin to reach initial cruise altitude. The cruise segment of the

mission itself, which is the longest component, occurs at a cruise step of 2,000 ft with a vertical

climb rate greater than 300 ft/min. after the cruise segment, the next phase is the descent to 10,000

ft at a descent Mach number. When the aircraft gets to 10,000 ft it initiates a braking manoeuvre to

slow down to 250 knots, and a secondary descent within the segment occurs which brings the

aircraft down to 1,500 ft. The penultimate phase is the approach and landing section which is stated

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to take around 5 minutes, which is equivalent fuel wise to a descent at mission landing weight from

20,000ft to 1,500 ft. The final segment is the taxi in which lasts 7 minutes.

In addition to the above mission critical segments, there has to be additional fuel to cover a

failed landing attempt, climb to 20,000 ft, extra cruise range of 200 nm, and a final approach and

landing. This requirement defined the fuel fractions shown in the next section of this report.

From the mission profile, and the house of quality a table was put together as an initial

guideline to any open ended velocities and design parameters as well as fixed requirements and

geometric limitations by ICAO codes or any other legal restrictions. This is shown on the figure below

Design Cruise Velocity 420 knots

Vertical climb rate during step cruise 2,000 ft/min

Wingspan <36m

Cabin average noise level 78 dB

Worst seat noise level 82 dB

Maximum take-off & landing distance

4327 ft

Figure 2.2.1 additional guidelines to design

3. Initial Weight Estimation-

3.1 Fuel Fractions

Figure 3.1.1 Fuel Fractions for mission segments

The figure above shows the output fuel fractions from the initial weight sizing module in the

AAA software. The warm up fuel fraction of 0.995 was used as a recommended value from AAA as

were the similar fractions of the taxi and take-off. The climb segment's fuel fraction was calculated

from the fuel consumption of 0.65 Ib/hr/Ib, the change in altitude from 1500 ft to 35000 ft at a rate

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of 2000ft per minute. Similarly the cruise segment fuel fraction was dictated by the fuel

consumption of 0.5 Ib/hr/Ib, a small change of 2000ft per minute, this time the initial cruise velocity

of 250 knots and a range of 1400 nm were also used on the calculation. The descent phases

consisted of same values of climb, and the landing phase was once again set as the AAA

recommended value of 0.9920. The extra climb, cruise, and descent section at the end of the mission

is to fulfil the requirement for extra load in the event of a 200 nm diversion.

3.2 Mission Segment Fuel Mass

Figure 3.2.1Fuel mass from fuel fractions

In the above figure the fuel fractions from 3.1 have been translated into weights in Ib's. AAA

was able to calculate these values for fuel used after the Take-off weight was found using the

regression line of a sample of aircraft. This method is shown in the next section 3.3 where the take-

off weight estimation is explained.

3.3 Take-Off Weight Estimation

The take-off weight estimation was done with the assistance of the AAA weight sizing

module. This allowed for the plot of a regression line with sample aircrafts empty and max take-off

weights to be compared with a line representing the actual conceptual aircraft weight estimations

from the fuel, empty, propulsion and payload weights.

On the next page on figure 3.3.1 is shown the plot created from the sample of 6 aircraft

similar in function and capabilities as the concept aircraft being developed. This line was used by the

Take-off weights sizing to find the coincident point between the regression line, and the conceptual

aircraft design line for a physical value pertaining to the empty and take-off weights.

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Figure 3.3.1 regression line from sample aircraft

As seen in the image below, figure 3.3.2 shows the design point intersection generated by

the weights module with the data added to the software for sample aircraft. Once these two values

for empty and take-off weight were worked out in order to get a more detailed weights list for

different components of the aircraft, weight fractions were used with another set of sample aircraft

from Roskam’s Design books (Volume V). This is shown in appendix C

Figure 3.3.2 take-off weight design point intersection with regression line

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Figure 3.3.3 Component weight table from class I weight sizing

The figure above was an output from comparing the weight fractions in class I weight

sizing module in AAA with sample aircraft. It allowed for the estimation for each component

weight, along with a value for the centre of gravity of the aircraft to be calculated from the

location in the geometry module of the software.

3.4 Useful Load for 28 PAX Mission Performance

Figure 3.4.1 estimated values for range for a number of different passenger configurations including

the required 28 passenger business class version

After the Take-off weight was estimated through initial and class I sizing, AAA allowed for

the estimation of approximate ranges at different passenger load capacities. As specified by the brief

the aircraft should be capable of a range of at least 3200 nm with a 28 passenger business class

configuration. On the above figure it can be seen that the actual value for the aircraft with the 28

passenger configuration is 4227.6 nm, which abides by the requirement. A figure of the

extrapolation is in appendix C.

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4. Aerodynamics-

4.1. Lift

After the weight sizing was completed, the aerodynamics module in AAA was implemented

to ensure that the aircraft would actually be able to lift itself and fly. This was used in conjunction

with the performance and geometry modules in an iterative process to gain the final geometric

configuration which directly impacted the lift generated due to shape, size and length of the wings,

empennage, fuselage and high lift devices.

4.1.1 Wing

The wing of the conceptual aircraft was generated first through the lift section of the

aerodynamic module to get first Cl max values for the aerofoil and then shaped in the geometry

module, before finally implementing its aspect ratio in the matching plot inputs on some of the

performance components. Through this cycle the current wing which is capable of holding the

aircraft level at cruise was sized. Below on figure 4.1.1.1 is a plot of the lift distribution on the

aerofoil when the aircraft is at a 4° angle of attack.

Figure 4.1.1.1 Lift distribution over wing at 4 degrees angle of attack

The data pertaining to the aerodynamics of the aerofoil was acquired with the use of XLFR

results of the lift and drag coefficients with regards to lift curve slope among other plots. This can be

seen in appendix D. Simple lift calculations from the condition above with a 4° angle of attack of the

aircraft show that the lift produced is equivalent to 130,168 Ibs , which is logical as when climbing

the aircraft needs to generate more lift than its own weight(115,402.1. Ibs). The calculation is listed

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in the aforementioned appendix section. At cruise, the wings are under a constant angle of incidence

of at least 1°, with more in the cruise climb condition, with a slight inclination.

4.1.2 Horizontal Tail

Figure 4.1.2.1 Lift distribution of tail at same condition of wing

Once the wing was finalized, the tail section was looked at. This sizing consisted of the

aerodynamics, geometry, and stability and control module. As such the tail was sized to provide the

required stability to the aircraft. The above plot shows the lift distribution under the same

conditions as figure 4.1.1.1. The lift curve slope with the non-linear region marked out for the

horizontal tail is shown in the appendix.

4.1.3 Vertical Tail

This was designed with a NACA 0012 aerofoil, symmetrical like the horizontal tail. As such

the potential to right the aircraft laterally comes from its sizing which was done in the stability and

control module of AAA.

4.1.4 Flaps and Slats

The high lift devices were sized alongside the matching plot, as it gave the maximum lift co-

efficient required for the wing at crucial take-off and landing segments of the mission profile. As

such the input and output variables used are shown in the appendix alongside a plot showing the

additional lift at various angles of deflection. The different types used on the wing vary from a

singular slat running along the leading edge, a split flap alongside the fuselage, and a double slotted

flap beside it. This configuration was used to get the optimal amount of extra lift from the devices on

deployment as figure 4.1.4.3 in the appendix shows the variation between the different types by

themselves with a drooped nose essentially emulating a deployed slat.

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4.2 Drag-

4.2.1 Drag Polars

The figure below shows the drag polars at all conditions of flight. These were acquired using

the aerodynamic module of AAA, implementing the class I drag calculations. Individual plots of all

the polars is listed in appendix E, along with the input values for the all polars plot.

Figure 4.2.1.1 all drag polars plot

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4.2.2 L/D Ratios

After calculating the lift and drag parameters of the wing, AAA allowed for the lift to drag

ratios for the most important mission parameters to be determined using the weights module.

These ratios heavily depend on the velocity of the aircraft as-well as the associated altitude of each

flight condition. On the next page on table 4.2.2.1 is shown the lift to drag ratios of both climbs and

both cruise phases in accordance with the project brief.

SEGMENT Velocity (Knots) L/D

Climb

250 18.62

Cruise

420 18.90

Diversion climb

250 17.76

Diversion cruise

420 13.22

Table 4.2.2.1 L/D ratios for important mission sequences

5. Performance-

AAA uses specific flight characteristics to generate a diagram to assist the aircraft designer to

choose the best optimized design point to work to. This “matching plot” displays the different

technical features and performance data on one graph, and where all of the points intersect, is the

optimal design point. Screen shots of all values input through the performance module of AAA are

listed in appendix F.

5.1 Components

5.1.1 Stall Speed

As the aircraft is meant to be a dual role platform the stall condition was calculated in the

event that the 28 passenger configuration runs out of fuel, and is gliding at the altitude where the

final descent would begin. This serves as a performance guideline in the event of emergencies. As

such the calculated weight ratio would be 0.7, and altitude is 1,500 ft. the lift coefficient values are

estimated values of aircraft wing performance based on the previous aerodynamics module. The

value of stall velocity for both clean and with flaps deployed was then deduced to obtain the

matching plot.

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5.1.2 Take-off Distance

The maximum take-off distance available is fixed at the length of the runway at London city

airport. As such it cannot exceed the specifications. This approach and take-off distance is limited to

a distance of 4327 ft on runway 09, and 4544 ft on runway 27 [3]. As such the take-off distance has

been set as 4100 ft, which then allows the aircraft to operate from either runway. The lift coefficient

value was set as maximum with flaps at 2.4 to emulate the deployment of high lift devices on take-

off.

5.1.3 Climb

The climb segment of the performance included a ratio of maximum thrust on take-off, as-

well as maximum continuous thrust for 8 seconds after take-off. Oswald factors at take-off were

noted down from suggested values presented by AAA.

5.1.4 Maximum Cruise speed

Maximum Cruise velocity was calculated in the performance module at the starting cruise

altitude of 35,000 ft. The design cruise speed (maximum) was set as 430 knots, with a weight

fraction of 0.9 to represent the aircraft mid cruise during an extended cruise mission profile. The

thrust to weight at cruise was estimated to be 0.175 to satisfy the convergence on the matching

plot. Oswald efficiency factor was chosen to be a suggested value from AAA

5.1.5 Manoeuvres/Turn

The manoeuvre segment of the performance module was linked to the instant where the

maximum cruise speed is evaluated, as such the altitude is 35,000 ft and the velocity at which the

aircraft is flying is 430 knots. The weight fraction at the instant of calculation is also the same at 0.9.

The load limit on the manoeuvre was set at 2.9 g. With a sudden manoeuvre of 2.9g at cruise it

would take a sudden change in power and course to cause the loads on the aircraft, as such the

thrust to weight of the condition was set at 0.52.

5.1.6 Landing Distance

The landing distance was calculated with instance at the end of the 100 PAX mission profile,

with a weight fraction of 0.85. Additional restrictions were the landing run available at London city

airport as in the take-off distance. In this case the landing roll was 2,400 ft, with a total of 4,100 ft for

the complete landing to occur. This is still below the smallest distance of runway 09 of 4,327 ft.

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5.2 Matching Plot

Figure 5.2.1 Matching Plot from AAA software including Legend

The matching plot is an important visual guide to a more in depth design point continued on

from the weight sizing. This allows for the key performance properties and data to be sized,

estimated and calculated to be translated into a physical object on to the geometry module of AAA.

Above it can be seen where all the lines intersect is the optimal point to which to design towards.

The plot allows for the Wing loading on take-off, thrust available, wing area, and thrust to weight at

take-off as seen in the figure below. It is created when defining characteristics of Stall, Take-off

distance, cruise, manoeuvre, and landing distance.

Figure 6.2.2 Data from Matching plot

point

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6. Geometry-

Figure 6.0.1 three view of aircraft in the geometry module of AAA 3.5

6.1 Wings & Empennage

The wings were shaped by initially by defining an aspect ratio and wingspan. After these two

properties were researched and chosen, AAA was used to build on, and mark out in detail all aspects

of the wing geometry, including the high lift devices implemented to maintain the maximum lift co-

efficient required by the matching plot for landing (2.4). Below is a figure displaying the as-well as all

the control surfaces marked out along it.

Figure 6.1.1 conceptual aircraft wing

Slat

Flaps

Aileron

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The fuel volume was calculated trough the class I sizing in the geometry module to ensure

that there was the capacity to carry the specified fuel load for the flight. AAA results showed that the

wing had the capacity to more than twice the mission fuel load with space for 53,959.2 Ibs on a 100

passenger configuration, and also had enough room for the 28 passenger configuration which lists

the mission fuel requirement for a 4,200 nm journey of 33,532 Ibs. The inputs and plots of the wing

and fuel calculations are listed in Appendix G.

Although from the figure 4.1.1.3 in appendix D it can be seen that the surface maximum lift

co-efficient of the clean wing is slightly below the required 1.6 from the matching plot, it is an issue

quickly rectified by altering the aerofoil which was chosen (NACA 1412) to add in slightly more

camber to increase the lift coefficient to the required state, this would entail the use of a custom

aerofoil using the base of the current one, mixed with components of another super critical type to

achieve the increase of lift required or add geometric twist to the wing to generate extra lift.

However the aerodynamic module results show the aircraft would fly with the current unaltered

wing.

A point to note is although on the geometry interface of the wing it shows the aerofoil as a

NACA 2412; it was the closest aerofoil in AAA’s directory which matched the 1412. The pylons

extending from the wing to the nacelles were chosen to consist of 0012 symmetrical aerofoils, to

provide extra lateral stability, as well as prevent any yawing moments which would come from lift

generated on one side due to camber.

AAA also enabled the control surfaces to be sized, from the ailerons, all the way through to

the elevator and rudder. These are also listed in the appendix along with their respective input and

output tables dictating sizes as-well as location of the hinge lines. The responsiveness of an

aileron/rudder/ elevator is the primary concern for a concept design in setting an initial value for the

control surfaces. Table 6.1.6 in the appendix was used as a first guess guideline to initially size these

with respect to chord lengths. It can be seen that the chord to chord ratios of the aileron, rudder and

elevator are 20%, 40% and 30% respectively. The span of these was decided by the span of the

sections around them, for example on the wing the aileron occupied the space that was left on the

outboard of the wing, and the rudder and elevator occupy most of their spans as-well.

6.2 Fuselage

There was a very specific set of guidelines for cabin comfort, as well as a variety of

differently dimensioned types of configurations within for passenger comfort and facilities during

flight. The figures below on the next page show the layout for the single economy class 100 PAX set-

up.

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Figure 6.2.1 Side view of cabin

From the side view a few features of the configuration are instantly noticeable. The spacing

between the seats is 12” which is 2” more than the minimum required. There is also a substantial

amount of legroom for the first row in the cabin, greater than the 18” required by the design

specification. There is also sufficient distance (greater than 5”) between the last row of seats and the

aft cabin wall to allow the seats to recline backwards. In the aft section of the fuselage as can be

seen from this angle there is one toilet and a one and a half galley unit next to the cabin wall. The

wing box underneath the cabin floor is marked out around the area where the wing intersects the

fuselage. In addition 3xLD3-45W Containers can be seen in the baggage area. There is 15 ft of space

between the cabin wall and the nose where the front emergency exits, flight attendants seats and

the cockpit will be located. Beneath this area on the nose, is the space dedicated to the permanent

flight systems. There is additional cargo space behind the wing box underneath the cabin, as well as

space for the APU at the aft section of the fuselage behind the toilet.

Figure 6.2.2 Top view of cabin with emergency exits marked out

From the top view of the cabin it is easier to see some aspects of the cabin configuration.

From here the Type A over wing emergency exit can be seen. Although only one is required due to

the dual exits on the back and front of the aircraft, if there would be a family created from this initial

concept, the simplification of one consistent over-wing fuselage section would be cheaper to

manufacture if it is the same throughout the different ranges of aircraft types available, so it is has

enough emergency exits to vacate an aircraft with more passengers than its current configuration

allows for. It also allows for additional safety the rare event that exits on either the forward or aft

sections of the fuselage are blocked. In such a case there would still be enough capacity to evacuate

the passengers quickly and safely. As per the regulation there is enough space for 5 and a half

trolleys worth of food trays, with three at the rear, and two and a half stacked towards the front of

the fuselage to feed 1.5x the passenger number (150).

Fixed equipment Wing box

Cockpit, front toilet, flight attendant

seats & forward emergency exits

Overhead carry-on baggage

compartment

Toilet Toilet

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Figure 6.2.3 Cross section of cabin

The cross sectional view of the cabin shown above shows the variation between the height

of an average adult whilst seated, the cross section of the baggage compartments, as well as the way

the rows are placed within the cabin. It has been set up as rows of 5 consisting of 3 and 2

components to ensure that the 100 passengers fit within the fuselage designed through AAA. The

height on the isle as well as under the overhead baggage compartments is in compliance with the

requirement in the brief of a minimum height of 6.89 ft on the isle, and a minimum standing height

of the passengers under baggage side bins of 5.5 ft. The AAA plot of the fuselage is listed in the

appendix under figure 6.2.4.

6.3 Landing Gear

The landing gear was placed underneath; offset 7.16 ft backwards from the aircraft

to the centre of gravity that was calculated through AAA as shown in the figure below. This

was to allow the front gear to share a small part of the load on the ground and on landing

conditions of stress.

Figure 6.3.1 location of landing gears

LD3-45W Container

Height of an

average

seated adult

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6.4 Ventral Fins

Ventral Fins were employed due to the distinct shape of the aft fuselage upsweep as well as

the low wing configuration.

Figure 6.4.1 view from underneath the aircraft of the aft section of fuselage

In Design Ventral fins are used when the upsweep angle of the fuselage at the rear is higher than

the standard practice range of 10-12 degrees. This would normally cause an increase in parasitic

drag from the fuselage caused by the separation of flow, however when the ventral fins are

employed on the back section, these reduce the drag as they prevent the premature separation that

causes the increase. They also are known to provide a small increase in the lateral stability of the

aircraft at a cruise condition [4].

7. Propulsion-

In AAA, the inlet area was the designing factor for the shaping of the nacelles, for instance

with the engine chosen the maximum fan diameter was 4.75 ft [5], so the nacelles were made to

have an inlet diameter of 6 ft. This was to allow space for the engine as well as allow for the

appropriate inlet area required from the propulsion module of 29.99 ft2 as seen on figure 7.1, and

the input for the nacelles geometry on figure 7.2 in appendix H.

Along with this the propulsion module allowed for the calculation of thrust required from

the drag of the aircraft. This was used to corroborate the data from the matching plot regarding

thrust required, which came out as 29,474 Ib from the propulsion module and can be seen on figure

7.3. To this effect engines required to produce the stated thrust available from the matching plot

were sourced out from similar sized aircraft. Subsequently a pair of GE Aviation model CF34-10

engines were chosen. These are the engines currently on the Embraer 195 series of aircraft and are

capable of producing between 18,820 – 20,360 Ib of thrust at sea level, which when put together

with a second engine comes close to but exceeds slightly the matching plot specified value of 32,942

lb of thrust available [5]. The AAA inputs and outputs of all calculations done in the propulsion

module are listed in appendix H.

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8. Stability and Control-

The stability of the aircraft is heavily linked to the location of the centre of gravity, this was

calculated on AAA’s class I weight module with regards to weight fractions of components as-well as

their location along the fuselage to provide point loads and ultimately dictate the centre of gravity

used in the stability module for sizing purposes.

8.1 Horizontal and Vertical Tail sizing

The horizontal and vertical tail was sized using the Class I sizing on the analysis section the

stability and control module of both longitudinally and directional to represent both horizontal tail

and vertical tail requirements. Below on figure 8.1.1 is seen the input tab for the horizontal tail.

Figure 8.1.1 Horizontal tail sizing

The vertical stabilizer has a similar input area shown below on figure 8.1.2. This is so that

directionally the aircraft is stable enough to fly straight and re-align itself in the event of a gust.

Figure 8.1.2 Vertical tail sizing

These values then updated the geometry automatically and implemented the correct sizing

to the aircraft model. It is important to size the horizontal and vertical tails so as to be able to

produce a controllable aircraft.

8.2 Static Margin & Neutral Point

After the tail sizing, the next task at hand was to check the stability of the aircraft with

respect to its neutral point and static margin. The static margin is usually represented as a

percentage and it represents the stability of an aircraft. When an aircraft has positive stability it

tends to return to its original flight path upon being perturbed by an external force such as a gust,

when an aircraft has negative stability, it means that when the aircraft is displaced by a gust, it

continues to further divert off from its original path. The higher the percentage the better, however

past a certain amount, too much stability results in an uncontrollable aircraft. Examples of static

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margins include the 747 with a static margin of 27%, the Cessna 172 with a static margin of 19%, and

the Lear jet 35 with a static margin of 13% [2]. Below on figure 8.2.1 is displayed the output value for

static margin acquired through the stability and control module of AAA which was calculated as

29.22%.

Figure 8.2.1 Stick free static margin

The input variables for the static margin calculation are listed in appendix I under figure

8.2.2. The neutral point location is a key variable as it is used later on in the design for a mode in

depth stability analysis with regards to distances from its location to the wing aerodynamic centre,

as-well as the centre of gravity [6]. It is used to create a c-g envelope for flight in the later stages of

design.

9. V-n Diagram-

Figure 9.1 V-n Diagram for Conceptual Aircraft from AAA Software

The V-n diagram is used to assess and visualise an aircraft’s flight envelope. It was generated

by using velocities from the matching plot modified to represent equivalent velocity at cruise

altitude, as-well as inputs which refer to the lift coefficient and drag coefficient values of clean

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configuration, landing and negative load factor, as-well as landing and positive load factor

respectively. These were used by AAA to draw up the flight envelope presented in figure 9.1. The

input and output parameters are listed in appendix J.

10. Conclusion-

The aircraft developed with the aid of AAA software is more than suited for the required

operations specified by the project brief, being able to operate from both of London city’s runway

configurations (09, and 27) as-well as having enough range to complete its inter-continental role in

its 28 PAX configuration. The aircraft is a suitable base for which to start a new family on. Its extra

considerations to layout with high capacity emergency exits over the wing allow the central body to

be kept the same in the event of any extensions to the fuselage to allow for a greater capacity. The

figure below illustrates this concept.

Figure 10.1 Proposed fuselage alterations for a 125 PAX configuration with the central section of the

fuselage remaining the same between the two models

When comparing the general properties of the concept aircraft shown on the table below to

the market analysis from table 1.2.1, the capabilities of the developed aircraft show promise with

good performance values with respect to range and passenger capability as it sits between the

ARJ21-700 and the larger B737-600.

Aircraft General Properties

MTOW Ibs

Wingspan ft

Range (nm)

PAX Capacity

Ceiling (ft) Fuselage Length (ft)

Regional 115,402.1

(52,345.5 kg)

116

(35.3 m)

1,436 (2,659 km)

100 41,000

(12,500 m)

100

(30.48 m) Inter-continental

4,227 (7,828 km)

28

Figure 10.2 Table of concept aircraft general properties

Although AAA has been invaluable to design, the software has its limitations. It is tailored to

Conceptual design only and follows key methods laid out in Jan Roskam’s Airplane design books

volumes I-VIII, there are a lot of assumptions in the values that can be defined in this particular

project, as well as software design limitations. For example, Oswald efficiency factor may only be

estimated along with other theoretical values in accordance with suggested values from AAA, as-well

as this the program can only calculate derivatives for the stability and control module at subsonic

velocities (up to Mach 0.7), and is not able to do so for transonic or supersonic flow [7].

Modular Central component

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Appendix A

Figure 1.2.2 graphs comparing values from table 1.2.1 as initial start guideline to the design

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Appendix B: Mission Profile

Figure 2.1.2 Table linked to mission profile requirements

Appendix C: Weights Estimation

Figure 3.3.4 Sample Aircraft for weights fraction calculations

Figure 3.4.2 Extrapolation plot showing design point of 100 PAX with the 28 PAX

configuration marked out by the convergence of blue marker lines.

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Appendix D: Aerodynamics

Figure 4.1.1.2 Aerofoil data on NACA 1412 XFLR Results at a Reynolds number of 1,000,000 [8]

Figure 4.1.1.3 CL max of the wing clean configuration

Calculation of lift of wing during 4° climb = 𝐿 =1

2𝜌𝑣2𝑆𝐶𝑙 =

1

2× 0.3 × (216(𝑚𝑠−1)2) × 140𝑚2 × 0.598 = 579,219𝑁 = 59,043 𝑘𝑔 = 130,000 𝐼𝑏𝑠

N.B values converted to 𝑚𝑠−1 from knots, and 𝑚2 from𝑓𝑡2, output in Newton’s divded by gravity

(9.81) then converted to Ibs from kg, 0.3= ρ at 35,000 ft.

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Figure 4.1.2.2 Lift curve from horizontal Tail displaying non-linear area

Figure 4.1.4.1 flaps and slats inputs and outputs on impact on lift

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Figure 4.1.4.2 Lift generated from high lift devices at various angles of deployment

Figure 4.1.4.3 Different types of high lift devices compared [6]

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Appendix E: Drag

Figure 4.2.1.2 Input and output values for the all polars plot

Figure 4.2.1.3 Drag Polar for take-off gear down condition

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Figure 4.2.1.4 Drag Polar for Take-off gear up condition

Figure 4.2.1.5 Drag Polar for landing gear up condition

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Figure 4.2.1.6 Drag polar for landing gear down condition

Figure 4.2.1.7 Drag polar for clean configuration

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Appendix F: Performance

Figure 5.2.3 Stall Speed Input tab of matching plot

Figure 5.2.4 Take-off distance input tab of matching plot

Figure 5.2.5 Climb input tab of matching plot

Figure 5.2.6 Maximum cruise speed input tab of matching plot

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Figure 5.2.7 Manoeuvre/turn input for matching plot

Figure 5.2.8 Landing distance input for matching plot

Appendix G: Geometry

Figure 6.1.2 Wing geometry dimensioned (top) with aileron dimensions (bottom)

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Figure 6.1.3 Fuel Volume wing storage calculation

Figure 6.1.4 Horizontal tail with elevator marked out and dimensioned with inputs and output values

displayed

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Figure 6.1.5 Vertical tail with rudder marked out and dimensioned with inputs and outputs shown

above the image.

Table 6.1.6 typical ranges of control surface flap chords and deflection angles [6]

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Figure 6.2.4 Fuselage Plot in AAA

Appendix H: Propulsion

Figure 7.1 Inlet Area Sizing

Figure 7.2 Nacelle geometry inputs

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Figure 7.3 calculation of Thrust from Drag

Appendix I: Stability and Control

Figure 8.2.2 Input values from stick free static margin

Appendix J: V-n Diagram

Figure 9.2 Input parameters for V-n diagram

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Bibliography

[1] A. T. A. Group, “Enviro.aero,” November 2010. [Online]. Available:

http://www.enviro.aero/Content/Upload/File/ATAG_EfficiencyGuide_REPRINT_web.pdf.

[Accessed 19 February 2014].

[2] D. C. Mares, ME 5539-ME 5309 Design and Analysis of Aircraft and Aerospace Vehicles Lecture

Notes, 2013-2014.

[3] L. C. Airport, “London City Airport> Airline partners> Facilities> Runway Facilities,” [Online].

Available:

http://www.londoncityairport.com/AboutAndCorporate/page/AirlinePartnersFacilities.

[Accessed 19 February 2014].

[4] D. P. Raymer, Aircraft Design: A conceptual Approach.

[5] G. Aviation, “GE Aviation>Engines>Commercial>CF34-10,” [Online]. Available:

http://www.geaviation.com/engines/commercial/cf34/cf34-10.html. [Accessed 20 February

2014].

[6] S. Gudmundsson, General Aviation Aircraft DesignApplied Methods and Procedures,

Butterworth-Heinemann, 2013.

[7] DarCorporation, “Advnaced Aircraft Analysis,” DarCorporation, [Online]. Available:

http://www.darcorp.com/Software/AAA/. [Accessed 23 February 2014].

[8] A. Tools, “Airfoil Tools,” [Online]. Available:

http://airfoiltools.com/airfoil/details?airfoil=naca1412-il. [Accessed 10 February 2014].