aiaa2009-example2

7/30/2019 AIAA2009-Example2

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AerospaceProject

Diogo MoreiraCarlos Henriques

Jos Afonso

Anabela Reis

Pedro Casau

Noel Leito

Simon Steidl

5558756285

56293

56323

56338

56341

66762


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1. Index2. The Design Philosophy................................................................................................. 5

2.1. Design Requirements ................................................................................................... 5

2.2. Mission Segments Overview .......................................................... ............................ 6

2.3. Competitive Analysis .................................................................................................... 7

2.3.1. Conventional Configuration ....................................................................................... 8

2.3.2. Box-wing Configuration............................................................................................... 9

2.3.3. Three-Surface Configuration - TSC........................................................................10

2.3.4. Blended Wing Body ................................................................ .....................................11

2.3.5. C-Wing Configuration ............................................................ .....................................12

2.3.6. Morphing Wing .............................................................................................................13

3. Concept generation .....................................................................................................15

3.1. Concept (1) .....................................................................................................................15

3.2. Concept (2) .....................................................................................................................16

3.3. Concept (3) .....................................................................................................................18

3.4. Concept (4) .....................................................................................................................20

3.5. Concept (5) .....................................................................................................................22

3.6. Concept (6) .....................................................................................................................23

3.7. Choice of Concept ......................................................... ................................................25

4. Maximum Take-Off Weight ......................................................................................27

6. Flight Envelope .............................................................................................................35

7. Wing design ......................................................... ...........................................................38

7.1.1. Wing Planform ..............................................................................................................38

7.1.2. Wing Airfoil ....................................................................................................................39

7.2. High Lift Devices ........................................................... ................................................44

7.2.1. Passive Lift Enhancement ............................................................... ..........................44

7.2.2. Active Lift Enhancement ...........................................................................................45

8.1. Landing Gear ....................................................... ...........................................................48

9. Engine Selection ...........................................................................................................49

9.1. Requirements and resulting tasks: .......................................................................49


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9.2. State of the art: .............................................................. ................................................50

9.3. Fuel Efficiency ...............................................................................................................51

9.3.1. Number of Engines ......................................................................................................51

9.3.2. Engine Type ......................................................... ...........................................................51

9.4. Noise Reduction ............................................................ ................................................53

9.5. Choice of engines .......................................................... ................................................54

10. Aircraft CAD Model ................................................................ ......................................57

10.1. Component Volume Estimation..............................................................................57

10.2. Component Weight estimation ............................................................. ..................58

10.3. Centre of Gravity ..........................................................................................................60

11. Aircraft Stability ...........................................................................................................61

11.1. Longitudinal Stability .................................................................................................61

11.2. Lateral Stability ............................................................. ................................................64

11.3. Control Surfaces Design.............................................................................................66

11.3.1. Rudder Design...........................................................................................................66

11.3.2. Aileron Design ........................................................... ................................................67

12. Structure Design ........................................................... ................................................70

12.1. Idealization .......................................................... ...........................................................70

12.2. 1stIteration .......................................................... ...........................................................70

12.2.1. Assumptions ..............................................................................................................70

12.2.2. Sketch ...........................................................................................................................71

12.2.3. Bending Moment ................................................................ ......................................72

12.2.4. Shear Strength ..........................................................................................................74

12.3. 2nd Iteration ......................................................... ...........................................................76

12.3.1. Assumptions ..............................................................................................................76

12.3.2. Sketch ...........................................................................................................................77

12.3.3. Bending Moment ................................................................ ......................................78

12.3.4. Shear Strength ..........................................................................................................80

13. Materials Selection ......................................................................................................82

13.1. Innovative typology ....................................................................................................82

13.2. Composite Materials ...................................................................................................82


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13.3. Natural Composites ................................................................ .....................................83

13.3.1. Sandwich Technology ............................................................................................84

13.3.2. Cost comparison.......................................................................................................85

13.4. Smart Materials ............................................................. ................................................85

14. Features of the designed airplane .........................................................................87

14.1. Take-Off and Landing distances .............................................................................87

14.2. Inputs analysis ..............................................................................................................89

14.3. Innovation ............................................................ ...........................................................92

14.4. Sustainability .................................................................................................................94

15.1. Production Costs ..........................................................................................................96

15.2. Operating Costs ............................................................. ................................................99

15.2.1. Crew ..............................................................................................................................99

15.2.2. Fuel ........................................................... .............................................................. .......99

15.2.3. Maintenance ..............................................................................................................99

15.2.4. Fees .................................................................... ........................................................ 100

15.2.5. Total Operating Costs ........................................................ .................................. 101

16. Conclusion............................................................ ........................................................ 102

17. References ........................................................... ........................................................ 105


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2. The Design PhilosophyAir transportation isnt affordable with current technology. Our planets limited

resources are pushing us to the edge of our creativity and its our role to come up

with new airplane designs.

We have been given the task to build a 150 passenger aircraft and these are its

most important characteristics: efficiency, ecology and sustainability.

Efficiency has to do with cost reduction. Making a more efficient design will

decrease fuel consumption while still making the airplane compliable with all

applicable regulations and purposes.

An ecological aircraft is one which produces less pollution. Current airplane

designs produce a lot of noise pollution and air pollution just during its flight.

Reducing the CO2 emissions into the atmosphere as well as noise is a major goal to

any new design.

Any new engineering entrepreneurship needs to focus on sustainability because

we have limited resources. We cannot keep digging Earths resources because their

renewal is not guaranteed. We need to think about recyclable and reusable

components. Having fewer resources also means higher costs thus having a

sustainable production loop assures well always have the required materials for

the job at an affordable price.

These are our main goals for the design we develop on the following sections.

2.1. Design RequirementsKeeping in mind the main goal of the project, our design should be able to satisfy

all of the imposed requirements. These several specifications are summarized in

the following table.

ID Description Value1 Capacity 150 passengers

2 Class Configuration Dual Class: (12 seats @ 36 pitch firstclass and 138 seats @ 32 pitcheconomy class)

3 Cargo capacity (bulk loaded) >7,5 ft3/passenger

4 Maximum payload capabilityFull single class 30 pitch passengercapacity (185 lbs/passenger) + full

cargo hold (8 lbs/ft3)

5 Maximum Range2800 nm with typical mission reserves

with full dual class passenger load,assuming 225 lbs/passenger.

6 Maximum Landing Weight (MLW)Maximum Zero Fuel Weight +

Reserves for Maximum Range Mission


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7 Typical mission (average) Ranges500 nm (50%), 1000 nm (40%), 2000

nm (10%)

8 Cruise speed Requirement0,78 Mach for Long Range Cruise

(LRC)

Objective: 0,80 Mach (LRC)9

Initial Cruise Altitude Capability atMTOW:

> 35,000 ISA + 15 C

10 Maximum operating altitude: 43,000 ft

11Maximum landing speed (at Maximum

Landing Weight):135 knots

12 Takeoff Field Length (TOFL), MTOW: 7000 ft (sea level), 86 F13 Community Noise ICAO Chapter 4 20 dB (cumulative)

14 Fuel Burn500 nm mission shall be requirement:< 41 lbs/seat. Objective: < 38 lbs/seat

15

airplane shall be certifiable to

appropriate FARs and entry intoservice

2018

16 Operating costs

8% or better (reduction); objective10%/seat or better operating cost

economics (Crew, Maintenance, Feesand Fuel at $2.50/US gal) than

current, comparably sized commercialtransports in typical US major airline

type operation.Table 1 - Mission Requirements.

2.2. Mission Segments OverviewThe mission segment can be described in Figure 1:

Figure 1 - Mission Segments


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The mission segments can be summarized in:

Warm up and Taxi;

Full thrust and Takeoff

Climb to cruise altitudeCruise at mach 0.8

Initial descent

Loiter

Final descent

Land

Climb to reroute altitude

Cruise to alternate

Descent to sea level

LandTable 2 - Mission segments

2.3. Competitive AnalysisThe purpose of this project is the design of a civil aircraft with 2800 nautical mile

range capability. Therefore, it makes all sense to watch closely our competition and

identify every potential new technology or development.

The main manufacturers of civil airplanes are Airbus and Boeing. Nevertheless, it is

important to investigate other manufacturers that might set newer developments

or trends, such as Bombardier.

Therefore, to carry out the preliminary sizing of our aircraft, it is important toconsider historical data of similar aircraft. The result of this study is presented in a

comparative table.

RequirementDescription

Objective A320-200Boeing

737 700Bombardier

CS300 ERA320 300ER

Year 2018 1998 1998 2013 2015

Number of seats 150 150 126 120 150

Maximum range

[nm]

28002600

3000

2885 2950 3000

Cruise speed [M] 0,8 0,78 0.781 0,78 0,8

Maximum Fuelburn for 500 nm

mission[lbs/seat]

38 58,1 52,3 n/a 41,7

Table 3 - Mission requirement objectives vs. current airplanes.


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From the data listed in Table 3 we notice the demand for a much more efficient

design, i.e. one that reduces dramatically the consumption per seat ratio while

keeping other parameters almost unchanged. The A320-300ER will enter service

in 2015 and its characteristics are fairly close to our objective.

Therefore, we consider the fuel burn requirement to be the most important design

driver because it demands great improvement in every aspect of the airplane, from

wing design to engine selection and so on.

The innovation in aircraft designs has reached a point where the current

configurations represent highly optimized design solutions. Therefore, the short

term available opportunities for innovations are in:

New products and technologies;

Process technology;

Technological innovations that present superior product substitutes.

On the other hand, the long term innovation is more ambitious and looks for new

designs that mitigate the fuel burn, emissions and noise. A lot of study has been

developed to provide an efficient solution for these problems.

So, in order to satisfy the imposed requirements, some concepts were evaluated.

2.3.1.Conventional ConfigurationThe conventional configuration, shown in Figure 2, has an obvious advantage of

being under study for the past years. Therefore, it has benefits in terms of design

development and post aircraft production.

Figure 2 - Conventional Aircraft Design.

The conventional design with a low horizontal tail is a common option since both

horizontal and vertical surfaces roots are attached directly to the fuselage. In this

configuration, the effectiveness of the vertical tail is large because the interference

with the fuselage and horizontal tail increase its effective aspect ratio. Large tail

areas are affected by the converging fuselage flow, however this can reduce the

local dynamic pressure.


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Winglets are used to improve the efficiency of the wing at the expense of some

extra weight. Their purpose is to reduce the aircrafts drag by altering the airflow

near the wingtips, which results in fuel savings. As a result, positive trade-off can

only be accomplished for longer than one hour flights1.

In Table 4 are summarized the main advantages of this concept.

Advantages Disadvantages

Design Less production cost;

Conventional Tail Minimum weight;Large effectiveness;

Converging fuselage flow;

Winglets More stability;Fuel savings;

Extra weight;

Table 4 - Advantages and Disadvantages of a conventional design.

2.3.2.Box-wing ConfigurationA visualization of the box-wing configuration can be found in Figure 3.

Figure 3 - Box-wing Configuration.

In a box-wing design, the tail horizontal stabilizer is extended and joined to the

wing. This is different from a joined wing, where the wings are connected to thevertical stabilizer. This configuration increases the overall span efficiency.

Nonetheless, the complexity of the structure leads to many concerns, because the

aircraft wings are under a large stress due to bending moments at the endplates

between the wings.

The engines are mounted on the tail of the aircraft below the upper wing, creating

a thrust line close to the CG, behind the passenger compartment. So, yaw control is

an important concern.1 From [6]


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Through the increased planform area, a larger amount of lift is generated than in a

conventional wing design. This larger area results in decreased take-off length,

allowing fuel-savings and lower stall speeds.


Increases the overall span efficiency Complexity

More lift Large stress at the endplates

Fuel Savings Yaw control

Shorter take-off and landing paths Production Costs

Decrease stall speedsTable 5 - Advantages and disadvantages of box-wing design.

2.3.3.Three-Surface Configuration - TSCThe three-surface concept intends to provide a higher lift to drag ratio (L/D) and

additional control surfaces on the aircraft. This design adds a canard to a

conventional aircraft, as can be seen in Figure 4.

Figure 4 - Three-surface Configuration.

Extra generated lift allows shorter take-off and landing distances and fuel-savings

or an increased range. The integrated canard stalls before the wing. This provides

the pilot enough time to react to the perturbation and recover. Furthermore,

integrated canard tends to move aerodynamic neutral point forward in the aircraft,

reducing the static margin, which decreases stability. The major disadvantages of

including a canard in a conventional configuration are a higher skin friction, higher

weight and lower stability.

The major characteristics of TSC are summarized in Table 6.


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Additional Lift Additional skin friction

Improved rotation behaviour Higher weightShorter take-off and landing paths Lower stability

Production Costs

Table 6 Major advantages and disadvantages of the TSC.

2.3.4.Blended Wing BodyBlended wing body is an alternative airframe design which incorporates new

design features and can be seen in Figure 5. This is not a new idea but only now,

with advances on material construction and computer-aided fly-by-wire, are its

huge gains in aerodynamic efficiency realistically achievable.

Figure 5 - Blended Wing Body.

It is highly fuel efficient due to the body extra lift. Aerodynamics of the overall

shape offer much lower drag, in part because it has no vertical tail. It is

complicated to control this concept due to the absence of tailfin. This issue can only

be addressed by a sophisticated computer flight-control system.


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Another concern about this concept is the passengers acceptance, many would be

travelling far from windows. Besides that, it is hard to evacuate so many people

from deep interior cabin in an emergency.


Higher produced lift Controllability

Lower fuel burn/emissions Passengers Acceptance

Lower Drag Difficulty in evacuation

Production Costs

Table 7 - Advantages and Disadvantages of BWB.

2.3.5.C-Wing Configuration

A concept using a C-Wing is visualized in Figure 6. The C-Wing design was

proposed as one way of addressing airport and manufacturing constraints. It

would also address the ineffective location of the outboard engine and the

excessive height of the vertical tail on a typical configuration. And possibly

improve the performance of the aircraft.

Figure 6 - C-wing Configuration.

In a C-Wing configuration the span can be reduced or the vortex drag can be

abridged at fixed span. The removal of the horizontal tail makes the use of aft-


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fuselage-mounted engines a possibility, eliminating some of the severe problems

with the original outboard engine location. Therefore, C-Wing is used as a primary

pitch control surface. The vertical and horizontal tip extensions offer an efficient

mean of fulfilling stability and control constraints . The horizontal C-Wing surfacesprovide more stability for a given area as they are not affected by the act fuselageflow field and are less affected by the wing downwash.

A C-Wing can be included in a three-surface configuration, providing a large

allowable range, with a relatively lightly loaded wing to simplify high-lift system

requirements and to accommodate passengers cabins in the wing.


Reduced Span Emergency

Reduced vertical tail height Aeroelastics control

Reduced wetted area Unconventional design

Table 8 - Advantages and disadvantages of C-wing configuration.

In spite of some pleasant features, the performance rewards for this design are not

significant, and probably not worth the risk associated with an unconventional

design.

2.3.6.Morphing WingThe design of fixed wing aircraft is constrained by the conflicting requirements of

several purposes. Devices such as flaps provide the present standard of adaptive

airfoil geometry, even though this solution conditions manoeuvrability and

efficiency, creating a design that is non-optimal in many flight regimes. Being able

to change the shape of the wings to reduce drag and power, which vary with flight

speed, would optimize fuel consumption so that commercial airplanes could fly

more efficiently. Morphing wings for flight control bring new challenges to the

design of control laws for flight. Because configuration changes move the

aerodynamic centre, control of the aircraft during planform morphing requires

attention.

Morphing structure is a challenge. A proper material for a morphing wing will have

to be elastic, flexible, resistant to different weather conditions, abrasions and

chemicals, have high recovery rates and a hardness number high enough to handle

the aircraft aerodynamic loads while in flight.

With current technology, these mechanisms are not practical because wingsweights would increase considerably as well as the cost-effectiveness. Smartmaterials instead, could be useful in the design of these new flight control devices.


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Shape Memory Alloys such as Nitinol are able to expand and contract with changesin temperature. Piezoelectrics and magnetostrictive materials such as Terfenol-Dare capable to do the same with the change in electric current or magnetic fieldrespectively. These smart materials would eliminate the weight problem and could

make the morphing mechanisms more practical.Studies are underway to investigate the availability of smart materials usage in a

wing that is able to adapt itself to any flight condition, gain maximum lift efficiency

and optimize its aerodynamic performance automatically. In Figure 7 can be

visualized an example of a morphing wing adaptable to different flight conditions.

Figure 7 - Flying efficiently at high speed requires small wings. Flying at slow speed for long

periods requires long wings. And with an asymmetric extension, morphing wings can provide

roll control.

Advantages Challenges

Improves aircraft performance Integrity of structures

Improve performance of conventional control surfaces Materials

Reduced drag Control systems

Improve Range

Reduce vibration

Table 9 - Advantages and Challenges of morphing wing.

While many questions remain unanswered regarding the utility of morphing air

vehicles, there is enough evidence to continue researching for performance

improvements with such mechanisms.


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3. Concept generationConceptual Design is regarded as the first step in designing a product. It is based

on the ideas or concepts that first emerged from the requirements and what the

product is intended for. The result is an outline or model that will be used later in

the development of the design.

3.1. Concept (1)The first concept highlights a V-tail mounted in the engine nozzle, winglets in the

wing tips and the engines mounting on the back of the fuselage, like what is shown

in Figure 8.

Figure 8 - Concept (1) with different engine inlets.

The advantages and disadvantages of the concepts main features are studied

bellow.

V-tail allows for a reduction in weight, it has less wetted area and thusproduces less drag. Although, a larger area is required for the same

performance and stability, so, structural reinforcement is needed.

Engine disposition on the fuselage allows for a reduction in noise. Also,there is no tail interaction with engine wake. Engine gets fuselage boundary


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layer (BL). Moreover, the engine displacement implies internal noise

amplification. Engine mounting place accounts for a heavier tail.

Blended Winglets allow improvements in stability due to their passivecontrol system which reduces the intensity of vortex. Winglets are

environmentally friendly. The noise is reduced in 6,5% 2 because of

degrading in needed thrust for take-off and landing. Blended Winglets can

save fuel cutting CO2 and NOx emissions by 3 to 6%3.Conventional design is an advantage in terms of design development and post

aircraft production.


Winglets

stability and

environmentallyfriendly Extra weight

Enginemounting

Noise reductionInternal noiseamplification

V-tail Less drag and weightStructural

reinforcement

Others Conventional design

Table 10 - Summary of advantages and disadvantages of concept 1.

3.2. Concept (2)The main difference between this concept and the previous one is the engine

mounting on top of the wings. Therefore, similar advantages and disadvantages

can be referred to the previous design concept for all characteristics except for the

engines.

2From [5]3 From [7]


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Figure 9 - Concept (2) with turbofan engines mounted on the top of the wings.

Figure 10 - Concept (2) with open rotor engines mounted on the top of the wings.

The main advantage of having the engine above the wing is to increase the laminar

flow, giving the wing a better efficiency. Also, engines on top of the wings create an

effect of noise shielding to ground observers.

The difference between the concepts 2 in Figure 9 and Figure 10 resides in the

engine, in Figure 9 the engine is a turbofan and in Figure 10 is an open rotor. Thebig advantage of open rotors over the turbofan engines is fuel saving. On the other

hand, open rotor produces more noise due to lack of physical encapsulation

nozzle which is also dangerous to the fuselage if one of the propellers cuts loose.


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WingletsStability and

environmentallyfriendly

Extra weight

Engine

Increases laminar flowNoise shielding

Open rotor is moreefficient

Open rotor producesmore noise and does nothave physical protection

V-tail Less drag and weight Structural reinforcement

Others Conventional design


3.3.

Concept (3)The next concept is known as a Blended Wing Body (BWB). Blended wing body isan alternative airframe design which incorporates new design features. This

concept produces a large increase in L/D due to fuselage contribution on lift and a

better load distribution. Also, BWB allows for less fuel burn, a higher passenger

and payload volume. However, BWB is more difficult to control due to small

moment arms to centre of gravity, therefore it is only controllable with winglets or

coupling with ailerons. Also, there is less structural stability for internal

pressurization. This is due to the sectional shape along the longitudinal length of

the BWB not being axisymmetric, which leads to difficulties in maintaininghomogenous pressurization inside the aircraft. Variations in pressure along cabin

can lead to spots with big differential pressure with the exterior, which in long-

term usage can lead to material and structural failure during the course of time

trough fatigue.

One of the biggest concerns with this concept is the emergency exits. A possibility

would be to include an emergency capsule to act in case of an emergency in order

to give passengers more time to flee.


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Figure 11 - Concept (3), Blended Wing Body.


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Wingletsstability and

environmentally friendlyextra weight

EngineIncreases laminar flow

Noise shielding

BlendedWing Body

Larger payloadEfficient high-lift wings

Increase of liftLess fuel consumption (10

to 25%)

Less stabilityLess structuralsuitability for

internalpressurization

Others New design

Unconventional

design associatedcosts


3.4. Concept (4)Concept 4 highlights the canard configuration, the mounting of the engines on top

of the wings, the usage of winglets and usage (Figure 13) or not(Figure 12) of a

vertical tail. As in concept (2), two concepts are presented, one with a turbofan

propulsion system (Figure 12) and another with an open rotor propulsion system

(Figure 13).

Figure 12 Concept (4) with turbofan engines without vertical tail


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Figure 13 Concept (4) with open rotor engines and a vertical tail

As referred in Concept (1 the introduction of blended winglets allows for an

increase in stability and a reduction in fuel consumption. In relation to engines,

advantages and disadvantages are already described in Concept (2), and may be

referred to that concept.Canard configuration allows for an improvement of stall characteristics,and has a big contribution towards improving stability of the overall

aircraft. Nevertheless it also introduces flow disturbances into the engine

inlet thus reducing its efficiency. Moreover, canards are known for their

complex sizing.

Tail omission accounts for less weight and structural stress on thestructure, but also a less manoeuvre capability. Although tail represents

more weight for the structure it also means an improvement in lateral

stability.

The disadvantages of not having a vertical tail (Figure 12) are taken in

consideration, so actuators will be present in the blended winglets, to account for

the loss of manoeuvre capability and lateral stability.

Conventional design as referred in Concept (1) will be advantageous in terms of

costs.


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WingletsStability and environmentally

friendlyExtra weight

Engine noiseshielding

Noise reductionOpen rotor as referred

Internal noise amplificationOpen rotor as referred

CanardImprovement of stall

characteristicsImproves stability

Complex sizingInlet engine flow

disturbances

TailA less weight

B improve lateral stability

A Less manoeuvrecapability

B more weight

Others Conventional designTable 13 - Summary of advantages and disadvantages of concept 4.

3.5. Concept (5)Concept 5 appears as bolder design. The main characteristics are the canard and

the engine mounting which have both already been described. The key of this

design is morphing wings. This kind of wing has better overall flying performance

and drag reduction. This comes from the fact that the wing adapts to the present

flight condition. In the present situation the aim of the morphing wing is to change

the wings aspect ratio providing better overall flight performance. However, a

morphing wing adds weight and structural complexity, as it requires a more

complex system to make the wing dynamic to the flight condition. A possible

solution to run the morphing wing system could be a truss system or a system

which uses smart materials (Smart Materials) as actuators. The truss system

would allow for the extension and retraction of the outer wing from inside the

inner wing. It would avoid the loss of interior space in the aircraft when the wing is

retracted by avoiding the need to have the same interior space as the extended

wings dimensions.


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Figure 14- Sketch of concept 5.

The type of fuselage allows the transportation of more payloads. However, this

comes with evacuation problems. Remembering the solutions showed for this

problem in the blended wing section, it can be solved with an emergency capsule.



characteristics

Improves stability


disturbances

Engine MountingNoise reduction due to

encapsulated engineInternal noise amplification

Morphing wingsBetter overall flying

performanceDrag reduction

Added weightStructural complexity

Others Larger PayloadAdditional loading andevacuation problems


3.6. Concept (6)The concept 6 is even bolder than the previous one. It has a lifting fuselage which is

a simpler structure that reduces wing overloading and makes it possible for an

easier accommodation. Round corners are a waste of space, but very useful from

structural point of view which help when dealing with cabin pressurization which

is the main problem with this design.


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Like it was seen previously, the engine mounting retards separation and it has a

noise shielding effect. The winglets produce lateral stability and improve wing

efficiency.

The influence of the stream coming from the canard, on the wing, will be reduced

due to their relative position. The wing is a low wing, and the stream comes from a

high canard, therefore the stream does not interfere as much as it would if the

canard was in line with the wing.

This design is also the safest amongst all other designs because lift distributes

itself more evenly throughout the airplane body reducing stresses and moments. It

is also known for its high lift over drag ratio and the ability to fly with one wing

only (if sized correctly).

Figure 15 - Sketch of concept 6 configuration A.


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Figure 16 - Sketch of concept 6 configuration B.



characteristicsImproves stability


disturbances

Engine Mounting

Noise reduction due to

encapsulated enginelatter separation

Internal noise amplification

Lifting FuselageReduces wing overload

Simpler structureEasy Accommodation

Difficult Pressurization

Others New design Evacuation problems

Figure 17 - Summary of advantages and disadvantages of concept 6.

3.7. Choice of ConceptFor the choice of concept a trade-off table was used. The idea of the following tableis to establish a standard trade-off to be used during the evaluation of design

concepts.

The criteria used to evaluate the concepts were stability, structural complexity,

environmental impact & sustainability and aerodynamic efficiency. Weight factors

were attributed to each one of the criteria according to their importance. The

values for the options range from 1 to 5, being the best option according to the

evaluation criteria 5 and 1 the worst case.


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Having established our criteria, the trade-off table illustrated bellow was built.

Options:Criteria:

Weightfactor

Stability 10 4 4 2 3 3 5

StructuralComplexity

25 3 4 2 5 1 4

EnvironmentalImpact &

Sustainability40 2 2 4 3 5 5

Aerodynamicalefficiency

25 2 3 5 4 5 5

Total scores 100 245 295 355 375 380 475

Table 15 - Trade-off table.

Looking upon the trade-off table, it comes to attention that concepts 4 and 6 arethe best choices. The concept 4 is a secure choice like it was said before, while the

concept 6 is bolder one and further research is needed therefore Concept 4 will be

our choice and Concept 6 might be studied later if we arent able to meet the

mission requirements.


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4. Maximum Take-Off WeightThe Maximum Take-Off Weight calculation described in [1] allows us to estimate

the fuel burn per seat ratio on the 500 NM mission, which has been identified as a

very important design driver requirement. We also test the maximum take-off

weights sensitivity to several parameters.

The calculations required input data is listed inTable 16 and a brief description of

each piece of data follows.

Input Data Value Source

Cruise Altitude [ft] 35 000 Requirements

Cruise Mach 0,8 Requirements

Cruise velocity 788 ft/s [1]Range 2800 Requirement 5

Thrust Specific Fuel Consumption0,38 Table Data4

Aspect Ratio 9,45 Airbus A3205

Structure Factor 0,5 Guess

Payload mass [lbs] 42 750 Requirements

Loiter Time 30Mission Segments

overview

Base Drag 0,008 Guess

Fuel Reserves + Fuel trapped [%] 6 RequirementsTable 16 - MTOW input data

Cruise Altitude Initial cruise altitude is taken from requirement number9;

Cruise Mach This is the objective Mach Number also listed inrequirement number 8;

Cruise Velocity This value is taken from the following equation whichdefines the cruise speed as a function of the Mach number and the cruise

height .

RangeMaximum aircrafts range according to requirements;

Thrust Specific Fuel Consumption Typical values for TSFC on a highBPR turbofan range from 0,05 on older equipment to 0,03 on new

experimental engines. For this first draft, the PWG turbofan engine data was

used.

4 http://www.jet-engine.net/civtfspec.html and also fromhttp://www.pw.utc.com/Products/Commercial/PurePower+PW1000G5 From [2]
http://www.jet-engine.net/civtfspec.htmlhttp://www.pw.utc.com/Products/Commercial/PurePower+PW1000Ghttp://www.pw.utc.com/Products/Commercial/PurePower+PW1000Ghttp://www.jet-engine.net/civtfspec.html


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Aspect Ratio For the first estimate, the Airbus A320 wing Aspect Ratiowas considered;

Structure Factor The first estimate was taken from Figure 2.5 in [1]

which provides the structure factor based on historical data;Payload Mass Passenger mass is 225 lbs (requirement 5) and passengercargo has, at most, 7,5 ft3 volume weighting 8 lbs per foot (requirement 4),

resulting in a total weight of 225+7,5x8 = 285 lbs per passenger.

Considering full dual class with 150 passengers the maximum payload is

150x285=42750 lbs;

Loiter Time This parameter is taken directly from the mission segmentdescription;

Parasitic Drag Equation 2.11 in [1] indicates that the parasitic dragcoefficient should be between 0,01 and 0,02. However we expect new

technology to reduce this parameter therefore was considered;

Fuel reserves The fuel reserves are indicated in the Mission SegmentOverview and should be 5% of flight fuel;

Trapped fuel 1% of flight fuel is considered to be trapped in the fueltanks (this value is taken from FAR Part 25 regulations).

With these assumptions and following the method described in [1] we get the

results listed in Table 17.

Maximum Take-Off Weight [lbs] 147 270Fuel Weight (2800 NM mission) [lbs] 30 886Fuel Burn on a 500 NM mission [lbs/seat] 59,5

Table 17 -1st iteration results

After the first iteration, the fuel burn is still far from our very demanding

requirement of 38 lbs per seat. However this wasnt unexpected because the input

data was based on historical values.

The identification of the maximum take-off weight sensitivity on the input

parameters provides us meaningful and useful information which may allow us to

specify some second level requirements.


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Figure 18 - MTOW vs Structure factor

Figure 19 - MTOW vs Thrust Specific Fuel Consumption.


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Figure 20 - MTOW vs Aspect Ratio.

In the previous set of figures we study the maximum take-off weight sensitivity to

a 10% deviation in the parameters from the value considered in Table 16. The

structure factor is by far the parameter which provides the highest change in the

calculations and it could be the key to a more efficient design.

A320-200 Boeing 737-700 A320-300ER

Structure Factor 0,55 0,49 0,43Aspect Ratio 9,48 9,45 9,8

Table 18 - Other airplanes design parameters

The trend is obvious. The structure factor of 0,5 considered before deviates from

current designs. One shall be able to achieve a lower structure factor.

Also, our wing design should have a higher aspect ratio in order to increase Lift

over Drag ratio with respect to current designs, but other solutions shall be

explored.

We consider new input values and recalculate the maximum take-off weight.

New Input Data Value

Aspect Ratio 12

Structure Factor 0,42Table 19 - New input values


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MissionSegment

InitialWeight

( )

FinalWeight

( )

WeightLoss

( )Source

Engine

Start-upand Take-

off

111 650 108 300 0,970 [1]

Climb andacceleration

108 300 104 840 0,968

Cruise todestination

104 840 94 510 0,902

Loiter 94 510 93 700 0,991

Landing 93 700 90 890 0,970 [1]Table 20 - Weight per mission segment summary

Maximum Take-Off Weight [lbs] 111 650Fuel Weight (2800 NM mission) [lbs] 22 008Fuel Burn on a 500 NM mission [lbs/seat] 37,9

Table 21 - 2nd iteration results

These new input values allow us to comply with mission requirement of 38 pounds

of fuel consumption per seat and should be considered as 2nd level requirements

during the remaining aircraft design.

5. Wing LoadingThe Wing Loading estimation requires the new input data which is presented in

Table 22 Summary of data input on wing loading estimation. and whose values

are explained below.

Input Data Value SourceOswalds efficiency

coefficient ( )0,8 Guess

Air density at sea level ( ) 1,225 Table Data

Table Data

0,3099 Table Data

0,9711 Table Data

Aspect Ratio (A) 12 Guess


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Where is the dynamic pressure at the beginning of cruise.

From the MTOW calculation we also have the weight at the beginning of cruise,

resulting in a wing area of:

This wing area achieves maximum range value as long as the assumptions are met.

One of the aircraft design constraints is the Take-Off Field Length because it is

limited to as stated in mission requirements. From this constraint and taking

into account the previously calculated wing area, were able to compute the

minimum Thrust required for take-off can be known by using the equations from

[1], which consider an historical based Take-Off Parameter TOP.

We get but we apply a safety factor of 2 because the aircraft must be

able to perform the take-off with one engine inoperative, therefore we get the

following result:

The pair is the so called Design Point which has to fit the set of constraints

imposed during each flight section, from take-off to landing. The set of constraints

was taken from [4] and then checked for design compliance with our design point.

Requirement Formula

Stall Speed

Take-Off

Cruise Speed

Landing


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Sustained Turn

Climb angle

Max Ceiling

Table 23 - Set of constraints from [4].

Figure 21 - Design Point and constraints.

Figure 21 shows that the design point verifies all the constraints. Furthermore,

with all the available data we are able to calculate the expected instantaneous turn

rate, from the formula in [4] considering cruise conditions.


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6. Flight EnvelopeFor a first structural analysis, a VN diagram on the point of view of project design

was made and for the calculations, some assumptions were made without knowing

yet the wing configuration.

So, for the positive and negative stall curves, we choose typical values for CLmax and

CLmin, meanwhile for CL we looked into linear airfoil theory and assumed a 2-D

infinite aspect ratio airfoil section with . For the positive and negative n

limits, we looked in FAR-25.


CL 2 Linear airfoil theoryCLmax 1,9 Design driverCLmin -1,4 Design driver

nlimit +2,5 and -1 FAR-25

Vdive 1,5xVcruise [1]Table 24 - Assumptions for Flight Envelope section.

Figure 22 - V-n Diagram.

Then, looking for gust loads in the normal direction in two different flight

conditions, we calculated the new load factors:

-Level flight:

Statistical gust load:

Response coefficient:


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Gust load velocity:

Lift before Gust:

Angle of attack before gust:

Angle of attack after load:

Load factor after gust:

-Dive Condition:

Statistical gust load:

Response coefficient:

Gust load velocity:

Lift before Gust:

Angle of attack before gust:

Angle of attack after load:

Load factor after gust:

For the calculations, we used statistical gust velocity values from [1], and also

equations (10.11), (10.12) and (10.17a). A summary is shown on Table 25.

Flight condition Altitude [ft] [ft/s] n

Level flight 35000 37,5 0,9559

Dive Condition 35000 18,75 0,7168

Table 25 - Summary table of calculations.

Figure 23 V-n Diagram with new load factors.


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Assuming a construction safety factor of 1.5 for the limits on load factors and a

quality factor of 1.15 to account with manufacturing defects, holes, connections,

etc, we got for the flight envelope in Figure 24

Figure 24 - V-n Diagram with safety and quality factor

Finally accounting for maximum gust loads, and assuring that the airplane stays in

the yield/elastic limit between the load factors n=-1 and n=3.5(gust loads - Figure

23), the flight envelope is depicted in the following figure.

Figure 25 Real flight envelope of the Project.


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7. Wing designThe most important goal during the wing design is requirements

compliance. As a result, our main driver was efficiency improvement in cruise.

Wing characteristics choice was focused on higher cruise performance, but with

reasonable performance on all other flight phases.

7.1.1.Wing PlanformDuring the design process, some assumptions were drawn regarding the wing

which are listed in Table 26.

Input Data Value SourceAspect Ratio (A) 12 Wing loading

Parasitic Drag ( ) 0,008 Wing loading

Oswalds efficiencycoefficient (e)

0,8 Wing loading

Planform area (S) 165 Wing loading

Dive speed (Vdive) 1,5Vcruise [1]Table 26 Wing Assumptions.

From the wing loading calculations, in order to maximize range, we achieve a

required planform area of: The high aspect ratios value decreases the

induced drag.

For the wing design several parameters need to be taken into account, and one of

the most important was the Cruise Mach velocity, M=0,8. With this value, we

looked for some historical data on [1], for , tapper ratio, and leading edge

sweep angle, .

Below, from historical data, weve got the following values:

Parameters Value Source

[1]

Tapper ration, [1]

Leading edge sweep angle, [1]

Table 27 - Geometrical Parameters from historical data.

With all of these geometrical parameters, the wing planform, should be similar to

Figure 26.


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Figure 26 - Dimensions of wing Planform [m].

7.1.2.Wing Airfoil

The lift coefficient for cruise was calculated assuming it is constant during the

whole cruise phase. With this assumption, the maximum ceiling at the cruise end

was approximately , that is less than the maximum ceiling of 43000ft from

requirements.

Looking for all the constrains, and doing some analysis on available airfoils, the

airfoil NACA 23012 is a good choice, designed for 0,3 lift coefficient, a little bitlarger than the expected design cruise lift coefficient.

Figure 27 NACA 23012 airfoil.

Pursuing cruise efficiency, drag needs to be reduced as much as possible. The

transonic airfoil design problem arises because we wish to limit or vanquish the

shock drag losses at a given transonic speed imposed by the requirements.


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Looking more carefully into the problem, we need to avoid the transonic wave drag

rise, characteristic of a drag divergence mach number.

From Korn equation applied to drag prediction on swept wings we verified the

drag divergence mach number, ,

7

Where, is an airfoil technology factor that can be assumed as 0,9 for this type of

airfoil. With previous assumptions, the obtained value was 0,89.

In order to avoid high mach numbers and shock waves for the present case, the

imposition of a leading edge sweep angle, , of changes the effective mach

velocity allowing shock waves to disappear.

From a compressible analysis for a time-marching Euler solver based on a

multidimensional upwind residual distribution method, Figure 28, we compare the

cruise conditions for with and without for an angle of attack ( ).

Figure 28 - Leading edge sweep angle illustration

We observe that the strong shock wave vanishes in the second condition.

Figure 29 - Mach distribution (=0 and

Mn=0.8).

Figure 30 - Mach distribution (=0 and

Mn=0.655).

7 From [8]


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Figure 31 - Grid illustration.

Running a batch analysis on[13] for different Reynolds, we got for the NACA 23012

a value of 0,0867/ ,Figure 32.

Figure 32 CL Vs plot.

From the Drag polar in Figure 33 it is possible observe that the minimum Cd is

around the design Cl cruise of 0,25:

-1,0

-0,5

0,0

0,5

1,0

1,5

2,0

-20,0 -10,0 0,0 10,0 20,0

Cl

CL vs

Re=60000

Re=110000

Re=160000

Re=210000

Re=260000

Re=310000

Re=360000

Re=410000

Re=460000


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Figure 33 - Drag Polar.

In order to test the finite wing, we defined a wing with the previously geometric

details:

Figure 34 -Finite Wing.

Root Chord[m] 6

Tip Chord[m] 0,9Span[m] 44,5

Area [m^2] 165Leading edge sweep [] 35

Aspect Ratio 12

Mean AerodynamicChord[m]

4,1

Mean GeometricChord[m]

3,45

t/c[%] 12Volume[m^3] 12,6Efficiency, e 0,89

Table 28 - Previous geometric assumptions.

Searching for the momentum reference location, we found a value for -

0,016 that coincides with the intersection point on Figure 35 and that is

obtained for the situation where .

-1,0

-0,5

0,0

0,5

1,0

1,5

2,0

0,00 0,05 0,10 0,15

Cl

Cd

Drag Polar

Re=60000

Re=110000

Re=160000

Re=210000

Re=310000

Re=360000

Re=410000

Re=460000


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Figure 35 - Cm Vs plot.

-0,5

-0,3

-0,1

0,1

-10,0 -5,0 0,0 5,0 10,0 15,0 20,0C

m

Cm vs


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7.2. High Lift DevicesThe High lift devices can be divided into two categories, the passive and active

devices. Depending on where they are positioned this devices are flaps (at the

trailing edge of the wing) or slats (at the leading edge of the wing). Both perform in

a way to achieve the necessary values for critical flight phases as take-off

and landing.

7.2.1.Passive Lift Enhancement

Trailing-edge devices:

Since the concept being followed is of a 150 seat passenger aircraft, and

considering similar aircrafts on the market, the most feasible flap to use is thesingle slotted flap. Considering the airfoil used for the wing, the same airfoil

NACA-23012 will also be used for the flaps. The considered airfoil will have a

chord of 20% of the wing chord. The following table exposes the flaps

characteristics:


Flaps deflection, [1]

Flaps cord ratio, [1]

Flaps area, m2 [1]

Table 29 - Flaps Characteristics.

The flaps area was calculated as follows:

Where, 0.8 is the ratio of flaps along the wing span.

When designing flaps the important solution that needs to be reached is the

increment in the maximum Lift Coefficient due to flaps introduction.

Where is the increment in the maximum lift for a flapped 2-D wing, is

the ratio of flapped wing platform area to wing area, and is an empirical

correction that accounts for wing sweep.

Another solution due to the introduction of flaps is the increment in the coefficient

of base drag of the wing.


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Where the coefficient is function of the ratio of the flap wing chords and the

coefficient is a function of the flap deflection.

Leading-edge devices:

For the slats the choice was the slotted leading-edge flap or more commonly

named Slat. This type of device is the equivalent to the trailing-edge flap. It works

by extending the leading-edge forward and downward, opening a slot and

increasing the wing section camber and area. This devices also generates an

increment in .

The final value for the is given by the sum of all increments due to high lift

devices and the of the wing without any high lift devices. It is as follows:

The value reached is x above the design value of 1.9 decided in the beginning of the

project. There is margin then to reduce the take-off distance.

7.2.2.Active Lift Enhancement

Active lift devices were not considered in this project due to its consequences in

terms of weight increment, complexity increment and for some cases, loss of

efficiency for the engines. The increment in would not suffice for the cost

that comes along with it. The trade-off was negative for these devices.


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8. Fuselage DesignThe fuselage design has a major role in any commercial airplane design because it

has to efficiently accommodate all the passengers and their cargo.

According to requirements listed in section 2.1:

The fuselage shall accommodate 12 first class seats and 138 second class

seats;

First class seats shall have a pitch of 36 inches;

Second class seats shall have a pitch of 32 inches;

Each passengers cargo shall have a volume of 7,5 cubic feet.

These sets of requirements provide the first design guidelines and all the fuselage

characteristics must not conflict with these constraints. Other dimensions are

taken out from historical data in [1] and are listed in inches for consistency with

the requirements.

Parameters Value [in]

1stclass seat width 22,5

1stclass aisle width 24

2nd class seat width 17

2nd

class aisle width 20Aisle Height 80

Headroom 65

Table 30 Summary of parameters took out from historical data in [1].

It is also necessary accommodate emergency exits and WCs, that are described by

FAR.


Emergency Exits 2-type I + 2-typeIII FAR

WC FAR

Table 31 - Emergency Exits and WC parameters according to FAR.

Two fuselage sections are provided in the following drawings:


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Figure 36- Fuselage top section.

Figure 37 - Fuselage cross section.

The fuselage empty volume must be enough to accommodate the cargo and fuel.

Some simple calculations are performed to assess this issue in Section 10.


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8.1. Landing GearThe fuselage will accommodate all the parts of the landing gear. It will be a

retractable tricycle type landing gear, with two-wheel bogeys at three points. This

is the optimal configuration to operate on paved runways, as it allows for ground

rotation both on landing and take-off without adding too much extra weight. The

main landing gear will be placed under the junction of the wing and the fuselage,

carrying approximately 90% of the weight while the other 10% will be carried by

the nose gear.

From ([1] eq 5.4) we estimate the main wheel dimensions:

From ([1], table5.10)

Main Wheel Diameter A=1,510 B=0,349 Diameter 51,8 in

Main Wheel Width A=0,715 B=0,312 Width 16,9 in

Nose Wheel Diameter A=1,510 B=0,349 Diameter 30,7 in

Nose Wheel Width A=0,715 B=0,312 Width 10.6 in

Table 32 - Main Wheel Diameter and Witdh ([1], table 5.10)

According to the results, the nose wheels can be modelled as 40% smaller than the

main wheels.

With these dimensions, the volume needed for the landing gear without accounting

for all the hydraulics, is approximately:

Using the method explained in section 11.1.5 of[1], it is possible to calculate an

approximate weight for both front and main landing gears.

Firstly, lets summarize input date and then present the results.

Parameter Value Justification gear

Kcb 1 Not a cross beam gear Main


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Kmp 1 Fixed gear Main

Ktpg 1 Not a tripod gear Main

Main landing gearlength 69.62 in Aircraft CAD model Main

Load factor 2.5 Flight Envelope Main

Number of mainwheels

4 Calculations Main

Number of main gearshock struts

2Estimation based on current landing gear

designs for similar aircraftMain

Stall velocity 100 kts Flight Envelope Main

Landing design grossweight

100646lbs

Landing with no more than half fuel weight Main

Knp 1 Fixed gear Front

Front landing gearlength

35.44 in Aircraft CAD model Front

Number of frontwheels

2 Calculations Front

Table 33 - Input data for landing gear weight estimation

As a result of calculations with this input data, landing gear weights where derived:

Total weight 1844 kg

Main gear 3714 kg

Nose gear 159 kg

Table 34 - Landing gear weights

Formulas for these previous calculations are omitted as they are length but simple.

They are simply comprised of several statistical coefficient times or powered toinput data provided and are easily obtained by referring to pages 261-262 of[1].

9. Engine Selection

9.1. Requirements and resulting tasks:


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Requirement Approach

Efficiency

Engine Type (Turbofan L-BP, Turbofan H-

BP, Prop-Fan)

Bypass RatioNumber of Engines

Supply of sufficient Power (Trust):

Static trust

Speed range

Engine Size

Number of Engines

Engine Type

Engine weight Engine Trust/Weight ratio

Noise Reduction

Placement

Insulation

Engine Type

Environmental requirements:

CO2 Emissions

Recycling ability

Residues due to maintenance

Engine Efficiency

Used Materials

Long Service intervals

Simple Design (small number of parts)

Maintenance costs Long Service Intervals

Reliability Simple Design (small number of parts)

Purchase priceSimple Design (small number of parts)

Common model/new developed

Table 35 Requirements and resulting tasks.

It is easy to see, that Requirements and Approaches are overlapping each other,

therefore a compromise must be found.

9.2. State of the art:Accomplished propulsions, focusing on currently used types of aircraft which have

similar characteristics to our project. Exclusively Turbo-fan engines are in use for

these aircraft types.


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Airplane Engine TypeTrust[kN]

Bypassratio

Airbus A319

(124-159 seats)

CFM International CFM56 82 - 151 Up to 6,5

International AeroEngines IAE V2500

98 - 147 4,5 5,4

Boeing 737-300/-400/-

700 (123-162 seats)CFM International CFM56 82 - 151 Up to 6,5

Bombardier CS 300

(120 145 seats)

Pratt & Whitney

PW1000G76 - 102 >>

Douglas DC 9 (all models,

90 - 172 seats)

CFM International CFM56 82 - 151 Up to 6,5

Rolls-Royce BR700 63 - 93 4 4,4

Pratt & Whitney JT8D 96,5 1,74

Table 36 - Comparison of aircraft engines.

Comparable Engines:

Engine Trust [kN] Bypass ratio

Pratt & Whitney PW6000 98 - 106 4,8 - 5

PowerJet SaM146 62 78 4,43

Iwtschenko Progress D-436 67 - 118 ~ 5

Rolls-Royce Tay 62 - 67 3,1

Table 37 - Comparable engines.

9.3. Fuel Efficiency9.3.1.Number of EnginesDue to a high Efficiency the number of Engines should be as low as possible.

Needs:

Sufficient amount of trust must be generated

The FAA requirements of minimum 2 engines must be fulfilled

One Engine must be enough for operating the aircraft

The minimum number of engines brings some more advantages:lightest solution

cheapest solution

less maintaining costs and time

For our requirements, most probably 2 engines will be the optimum.

9.3.2.Engine TypeDue to the useful flight Mach number, the focus is on Prop-fan engines and on

High-Bypass Turbo-fan engines.


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High Bypass Turbo-fan:A high bypass ratio gives a lower (actual) exhaust speed. This reduces the specific

fuel consumption, but reduces the top speed and gives a heavier engine. A lower

bypass ratio gives a higher exhaust speed, which is needed to sustain higher,usually supersonic, airspeeds. This increases the specific fuel consumption. Bypass

ratios of modern engines are range between 5 and 11 (Rolls Royce Trent 1000).Advantages:

Quieter around 10 to 20 percent more than the turbojet engine due togreater mass flow and lower total exhaust speed.

More efficient for a useful range of subsonic airspeeds for same reason,cooler exhaust temperature.

Less noisy and exhibit much better efficiency than low bypass turbofans.

Disadvantages:Greater complexity (additional ducting, usually multiple shafts) and theneed to contain heavy blades.

Fan diameter can be extremely large, especially in high bypass turbofanssuch as the GE90.

More subject to FOD (Foreign object damage) and ice damage. Top speed is limited due to the potential for shockwaves to damage engine.

Thrust lapse at higher speeds, which necessitates huge diameters andintroduces additional drag.

More Space is required: More ground clearance at under wing assembly, no

possibility of integration into the fuselage (noise reduction).

Geared TurbofanAs bypass ratio increases, the mean radius ratio of the fan and LP turbine

increases. Consequently, if the fan is to rotate at its optimum blade speed the LP

turbine blading will spin slowly, so additional LPT stages will be required, to

extract sufficient energy to drive the fan. Introducing a (planetary) reduction

gearbox, with a suitable gear ratio, between the LP shaft and the fan, enables both

the fan and LP turbine to operate at their optimum speeds. Typical of this

configuration are the long-established Honeywell TFE731 (already introduced in

1972), and the recent Pratt & Whitney PW1000G.

Prop Fan/Unducted FanAn Unducted Fan or Prop-Fan is a modified turbofan engine, with the fan placed

outside of the engine nacelle on the same axis as the compressor blades. The

bypass ratio of prop fan engines are around 20:1.

Advantages:Higher fuel efficiency
http://en.wikipedia.org/wiki/Specific_fuel_consumptionhttp://en.wikipedia.org/wiki/Specific_fuel_consumptionhttp://en.wikipedia.org/wiki/Specific_fuel_consumptionhttp://en.wikipedia.org/wiki/Mass_flow_ratehttp://en.wikipedia.org/wiki/GE90http://en.wikipedia.org/wiki/FODhttp://en.wikipedia.org/wiki/FODhttp://en.wikipedia.org/wiki/GE90http://en.wikipedia.org/wiki/Mass_flow_ratehttp://en.wikipedia.org/wiki/Specific_fuel_consumptionhttp://en.wikipedia.org/wiki/Specific_fuel_consumptionhttp://en.wikipedia.org/wiki/Specific_fuel_consumption


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Potentially less noisy than turbofans.

Could lead to higher-speed commercial aircraft, popular in the 1980s duringfuel shortages

Disadvantages:typically more noisy than turbofans

complexity

cruising speed limited because of blade tip speed

Regarding the terms of fuel efficiency, a Prop Fan engine seems to be a possiblesolution. A further approximation between Turbo Fan and Prop Fan is conceivable.According to our requirements regarding to cruise speed and noise reduction,most probably a high bypass turbo fan engine will be used.

9.4.

Noise ReductionThe noise, produced by the engine(s) depends on the following factors:Engine Type

Insulation

PlacementThe engine type is rather defined by other considerations like efficiency and cruisespeed requirements. In theory, prop fans could be more silent than turbo fans,especially during the approach. The approach is after the take off, the second mostimportant part of the flight mission in regards to noise emissions. Actually existingprop fans in the past (1980s) produced higher noise emissions than turbo fan

engines, especially inside the cabin.For a given engine type, different approaches for noise reduction can be found:

Fan blade geometry

Exhaust duct covers whose edges are serrated in a toothed pattern to allowa quieter mixing of exhaust and outside air (chevron/serrated nozzles)

Insulation:Insulation in case of pylon mounted engines is barely possible. As mentioned apossibility is to place sound absorbing materials inside of the air inlet and the fanair-ducting.

Placement:A much more effective way for noise reduction is provided by the engineplacement. The problem with the actual fuselage designs is, that the commonlyused turbofan engines are too large in diameter to be integrated in the fuselage.

Possible solutions:additional air-ducting with sound absorbing materials at the pylons

partly integrated into fuselage

over wing position: The wing can reflect a certain amount of the noise

mounting above the tail

new fuselage design with completely integrated engines


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9.5. Choice of engines

From our Design Point:

T/W = 0,22

T = 109 KN

Number of engines = 2

Tmin = 54,5 KN/Engine

We can consider two approaches Use of an already existing engine or

development of a completely new engine. To use an existing engine has following

advantages:

- No development costs

-

Availability, no developing period- Proved concept. Experiences about reliability, operating costs etc.

But this solution has the disadvantage that it is difficult to meet all requirements as

thrust, fuel consumption etc.

Available Engines which meet approximately our requirements:

Thrust[kN]

Weight

[kg]

Power/WeightRatio

[Nt/Nw]

Spec. FuelCon-

sumptionlb/(h*lbf)

1strun

Comment

CFMLeap X

81 - 157 - - - 2012Very high BP-ratio, high

efficiency

PW1000G 62 - 104~

1750- - 2007

Geared fan, very high

BP-R, high efficiency

PowerJetSaM146

62 - 78 - - 0,63 2006Low maintenance costs,

silent, efficient

PW6000 82 - 109~

2450~ 4,4 - 2000

Simple design, low

maintenance costs, low

fuel consumption

GE CF34-8 62 ~1120

~ 5,5 0,68 1999 High BP-R, well field-tested

RollsRoyceBR700

63 - 95 - - - 1994

Well field-tested, low

noise and pollutant

emissions

ProgressD-436

62 - 921400 -

1600

~ 4,4

5,80,63 1993

Efficient, high power to

weight ratio.

IAE V2500 98 - 1462240 -

2540

~ 4,3

5,7- 1988

High BP-R, high Power

to Weight ratio


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RollsRoyce Tay

62 - 69 - 4,2 - 1984Low BP-R, low Power to

Weight ratio

CFM 56 82 - 1511940 -

4000

~ 3,7

4,2

~ 0,55

0,651974

High BP-Ratio, well

field-tested

From the listed engines following engines are most suitable:

CFM International CFM Leap XThe CFM Leap X is still in development. It is supposed to achieve its fist run in

2012. The Leap X is an evolution of the very well established engine family CFM 56.

Compared to the current CFM 56 models it should have 16% lower fuel

consumption. The Bypass Ratio is very high, around 10:1.

Pratt & Whitney PW1000GThe PW1000G is a geared fan engine. This means that the fan and the turbine,

connected with a gear (ratio ~ 3:1), can work at their optimal speed. This increases

the efficiency of both modules and decreases as well noise emissions and fuel

consumption. The PW1000G is designed to have lower manufacturing costs as

current engines.

Figure 38 - Manufacturer's data.

Additional Data which is not available from the manufacturer. These values are

estimations, based on comparison of engines with similar thrust range, BP-ratio,etc.

Property Value Unit Source

Weight 2750 kg Estimation

Overall Length 3000 mm Estimation

Overall diameter 2000 mm Estimation

TSFC 0,38 Estimation


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10.Aircraft CAD ModelHaving defined important aircraft geometry such as the wing and fuselage, we are

able to draw a CAD model which includes also lifting surfaces included in the

concept generation process such as the canard and the winglets.

This aircraft model allows for improved geometry visualization which helps the

definition of important relations between the several elements that build the

aircraft. These relations include:

Wing root placement along the fuselage;

Canard placement with respect to the wing;

Landing gear placement;

Payload placement.

The CAD aircraft model geometry is controlled by means of a Design Table (in

Microsoft Office Excel) which easily handles changes in the design. This feature

allowed a very close interaction with the stability calculations which ultimately led

to the aircraft model in Figure 39.

Figure 39 - Aircraft CAD Model

10.1. Component Volume EstimationThe aircrafts CAD model is used for empty volume estimation so as to comply with

storage space requirements. Firstly we compute fuel and cargo required volume

and those results are cross-checked with CAD geometry output.


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Jet fuel density ( ) Table Data [2]

Jet fuel mass ( MTOW calculation

Cargo Volume RequirementTable 38 - Occupied volume calculation input data

From the input data one may easily compute the fuel volume.

From the geometric data we are able to compute the empty volumes. The fuel is

stored in the wings whilst the cargo is stored in the fuselage aft section.

Figure 40 - Cargo (red) and Fuel (blue) placement

Output Data Value

Cargo empty volume

Fuel empty volumeTable 39 - Fuselage empty volumes

From the data in Table 39 we conclude theres enough empty volume to

accommodate cargo, fuel and other airplane equipment.

10.2. Component Weight estimationThe component weight estimation which led to the inertia properties calculation

within the Solidworks software was performed according to the followingprocedure:

1. Draw airplane geometry;

2. Insert dummy components within the fuselage compartments which

simulate the passengers and cargo weights;

3. Insert fuel within the wing shell;

4. Assign density properties to the components materials in order to be

compliant with the calculations performed in section 4

The results obtained with this estimation method are listed in Table 40.


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Component Weight Source

Wing 4934 kg Estimation

Fuselage 10 107 kg Estimation

Canard 143 kg Estimation

Winglet 77 kg EstimationNose 113 kg Estimation

Back 78 kg Estimation

Engines 3500 kg Estimation8

Landing gear 1844 kg Section 8.1 of this report

Fuel 9983 kg Section 10.1

Passengers 15 309 kg Requirement 5

Cargo 4082 kg Requirement 1,3 and 4Table 40 - Component weights

The winglets weight might seem too high when compared to other surfaces

weights but this high estimate is expected to account for additional structural

reinforcement to hold the winglets and the rudder mechanism in place. On the

other hand the canard has a low weight value because it is mainly a trim and

control surface which does not have high loads.

In addition, besides the method described above, the results where compared to

those obtained using the method described in chapter 11.1 of [1]. Although this

method stands on statistical data as well as a number of parameters that depend

on the type of system used, its accuracy is limited due to the non-inclusion of the

most state-of-the-art technological innovations as well as the lack of consideration

for interior furbishing, seating, galleys, cargo holding structures and so on.Therefore, systems like fuselage or lifting surfaces are calculated weighing less

than with SolidWorks. The components weights obtained with [1] are presented

on Table 41 - Alternate component weights

Component Weight

Wing 5445 kg

Fuselage 5438 kg

Canard 102 kg

Landing gear9 1844 kg

Table 41 - Alternate component weights

The cargo location can be adjusted during the design process if any change in the

centre of gravity should be necessary.

8 There is little information on this subject other than some P&W marketing articles or speculationfound on the internet. As a result these values are a rough estimation.9 Landing gear weights only source is the Corke method, thats why the value is the same in the twotables


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10.3. Centre of GravityThe aircrafts centre of gravity (CG) is easily computed using the CAD software and

is depicted in Figure 41.

Figure 41 - Aircraft centre of mass

In order to compute the aircraft stability characteristics the CG distance to theleading edge is of great importance and we are able to determine it from data in

the CAD software.

The orthogonal reference frame depicted on the CG in Figure 41 is the Principal

Axis of Inertia and has the axis along the aircrafts length and the to the

zenith with a slight angular displacement. The axis is normal to the symmetry

plane. The principal inertia moments are listed below.

The full geometry design process which ultimately led to these values is explained

in the next section.


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11.Aircraft StabilityStability describes a systems response to any perturbation on its equilibrium

state. The system is stable if it returns to the given equilibrium point and is

unstable otherwise. In order to prove the aircrafts stability the nomenclature and

formulas described in [3] are used. A longitudinal stability study is presented

followed by a lateral stability assessment.

11.1. Longitudinal Stability

Input Data Value SourceDistance between canard

a.c. and wings a.c. ( ) 20 m GuessWings mean chord ( ) 4,272 m Wing Design

Wings area (S) 164,74 m2 Design Point

4,97 rad-1 Wing Design

4,97 rad-1 Wing Design

-0,016 Wing Design

496 813,599 N MTOW calculation

0,8359 Table Date

53,4226 RequirementsTable 42 Input data for longitudinal stability section.

We study the stability of a canard/wing configuration and first of all some

important coefficients and formulas are introduced.

is the airplane neutral point, i.e. the center of gravity location (with respect to

wings leading edge) which provides neutral stability and is the horizontal

volume coefficient. As a preliminary analysis weve neglected the downwash

contribution as well as the propulsive systems influence.


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These are the dimensionless moment equations, where the coefficient of lift is

given by the following equation whose coefficients were already defined except

which is the wings angle of attack and which is the canard incidence angle

( ):

In order to have a stable aircraft the following conditions need to be met:

The positive is easily achieved with a canard, provided it has enough incidence

angle. Having the center of gravity before the airplanes neutral point provides

negative . is an unknown variable but it must be within an acceptable range.

must lie within a prescribed range as stated in [1]:

The longitudinal stability analysis provides the canard dimensioning. During thewhole analysis we shall consider:

The canard and the wing are identical (same aerodynamic coefficients but

different surface values);

The canard is a stabilator, i.e. it changes its incidence angle in order to

trim the airplane, requiring no flap mechanism;

The take-off is performed at 6000 ft which is above the highest airport

runway on the USA;

The canard is near stall during take-off (i.e. );

At take-off the airplanes speed must be 20% higher than the stall speed,

thus were able to compute the target :

The trimmed angle of attack is found solving the equation .


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With the iterative method described in Figure 42, we are able to find the canard

area and incidence angle during take-off for any valid CG location which is given

with CAD software.

The coupling between the aircrafts geometry and the longitudinal stability

algorithm requires another iterative process which is described in Figure 43. The

process stops whenever the difference between two iterations is below a certain

tolerance. After four iterations the process converges and the obtained values arelisted in

Figure 43 - Stability design process

Updatedesign table

Updategeometry

Calculate CGlocation

Run stabilityalgorithm

Figure 42 -Representative diagram of iterative method to find the canard area


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0 29,7 0,709 - - - - -1 22,9 0,712 -1,2 5,66 0,302 14,4 0,212 22,9 0,712 -1,19 5,66 0,302 14,5 0,213 22,9 0,712 -1,19 5,66 0,302 14,5 0,21

Table 43 - Iteration results

Regarding the stability characteristics we notice that the pitch stiffness is within

the acceptable value range leaning towards the lower end, which means increased

stability. The trimming is accomplished at but this is not the geometric

angle of attack and corrections are made in order to account for the airfoil curvature.

Having the canard near stall is, despite what it seems a safe solution because the

canard loses lift before the wing and therefore reduces the angle of attack,preventing it from ever entering stall.

11.2. Lateral StabilityLateral stability is assessed mainly by two coefficients:

- rolling moment coefficient;

yawing moment coefficient.

More precisely, were interested in their derivatives with respect to the sidesl ip

angle . According to [1] and its formalism, the lateral stability criterion is

specified below.

This data is hard to determine, requiring most of the times wind tunnel

measurements. As a first approximation we consider as suggested in

[1] and is determined from empirical expressions.

Because our design doesnt have a vertical stabilizer we consider the winglets

acting as vertical stabilizers, but first we estimate the fuselage contribution to .

The input data in Table 44 is used.

Input Data Value SourceFuselage volume ( ) 333,8 m3 Fuselage Design

Fuselage height ( ) 3,55 m Fuselage DesignFuselage width ( ) 3,55 m Fuselage Design

Wing sweep angle ( 35 deg Wing Design


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Aircrafts Neutral Point ( 0,922 Longitudinal stabilityAircrafts CG ( ) 0,712 Longitudinal stability

Wings Neutral Point ( ) 1,52 Wing Design

Winglets neutral point

( ) 0,25 NACA0012 dataWinglets chord ( ) 0,9 m Wing Design

4,97 Wing Design

Table 44 -

aiaa2009-example2

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