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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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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
IdeasDesign 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
2
𝜌𝑣2
𝑆𝐶𝑙 =
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 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 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].
