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Chapter 1INTODUCTION TO AIRCRAFT DESIGN

1.1 Understanding Design

Microsoft Encarta Encyclopedia explains “design” as:

“…creating an object's form and function. Design can involve making products, machines, and structures that serve their intended purpose and are pleasing to the eye as well.”

Encyclopedia Britannica’s Dictionary provides numerous meanings of “design” as a noun as well as a verb. Most of them are given as under:

As a noun As a verb a particular purpose held in view by

an individual or group deliberate purposive planning a mental project or scheme in which

means to an end are laid down a deliberate undercover project or

scheme a preliminary sketch or outline

showing the main features of something to be executed

an underlying scheme that governs functioning, developing, or unfolding

a plan or protocol for carrying out or accomplishing something (as a scientific experiment); also the process of preparing this

the arrangement of elements or details in a product or work of art

the creative art of executing aesthetic or functional designs intention, plan

to create, fashion, execute, or construct according to plan

to conceive and plan out in the mind to have as a purpose to devise for a specific function or

end

In the light of the above description, we can now understand design as an activity involving the setting out of a strategy that ends up as a solution to the given problem which at the same time remains within the existing constraints. In engineering, the design process is of fundamental importance which determines the manufacturing scheme of a product. Any shortcoming or defect in a product’s field service is finally attributed to its designer.

1.2 Initiation of a New Design Process

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Innovation and creativity are continuously required in the technological dynamics of any dimension of applied knowledge. The reasons for the initiation of a new design process are listed as follows:

Customer requirement New requirement based on future market trends or research New technology or innovation

1.3 Aircraft Design

“Aircraft design is both an art and a science.”John D. Anderson, Jr. Aircraft Performance and Design McGraw–Hill, 1999, pp. 381

An aircraft is a very diverse system whose production demands input from the various subfields of aeronautics and avionics. Raymer declares aircraft design a separate discipline of aeronautical engineering, apart from aerodynamics, structures, propulsion and controls. An aircraft’s design is the outcome of an iterative process that is accomplished through the optimization of various candidate configurations and combinations. The following chart shows the methodology of the conceptual design process of an aircraft.

Chapter 2

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Design Requirements / Specifications

Concept

Analysis

Sizing and Trade Studies

Optimized Result

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REQUEST FOR PROPOSAL AND MISSION PROFILE

RFP, or Request for Proposal, is a proposal made by a commercial organization inviting bids from possible suppliers of a product or service, or by a government or other funding agency inviting bids from possible research bodies. The RFP given to me was for the AIAA Foundation Undergraduate Team Aircraft Design Competition, 2008 – 2009.

The RFP demanded for conceptual design of a new commercial transport aircraft design with a capacity for 150 passengers in a dual class configuration. Airbus A320 and Boeing 737–600 were declared as the existing comparables. The RFP required the aircraft to be environmentally friendly, compatible with the existing airports’ infrastructure and fuel efficiency. Also, it constrained the use of turboprops. Other technical specifications that RFP had asked for are listed as follows:

1. Passenger weight: 185 lbf2. Baggage: 60 lbf/paxx 3. Maximum Range: 2800 n mi 4. Cruise: 0.8 M5. Maximum Operating Altitude: 43,000 ft (Absolute Ceiling)6. Maximum landing speed (at Maximum Landing Weight): 135 knots 7. Takeoff Field Length (TOFL), MTOW: 7000 ft

2.1 Mission Profile

With the specifications at hand, I started the design process. A mission profile based on the designer’s choice was carved out, displayed as follows:

The mission profile is stationed as follows:

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0 – 1 Taxi and TO1 – 2 Climb2 – 3 Cruise: 2600 n mi @ 36,000 ft, 0.8 M3 – 4 Descent4 – 5 Loiter: 20 min5 – 6 – 7 Attempt to Land7 – 8 Divert (climb)8 – 9 Cruise: 200 n mi @ 20,0009 – 10 Descent10 – 11 Loiter: 20 min11 – 12 Landing

The mission profile ensures that its accomplishment would mean the satisfaction of most of the requirements.

Chapter 3

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CONCEPT SELECTION

An artistic sketch of the aircraft was made, after selection of the choices from a morphological matrix.

3.1 Morphological Matrix

The morphological matrix was formed as shown:

ChoicesAircraft

ConfigurationConventional Span

LoadedFlying Wing

Tandem Wing

Multi Fuselage

Flatbed

No. of Fuselages

1 2 3

Wing Configuration

High Wing Mid Wing Low Wing

Wing Sweep Forward BackwardTail Type Conventional T–Tail H–Tail V–Tail

No. of Tails 1 2 3No. of

Engines2 3 4 5 6

Wing Twist Aerodynamic GeometricSweep Type Fixed Variable Double TripleWing Position Anhedral Dihedral

Based on the selections of this morphological matrix, the following layout was made for the as–drawn aircraft:

Fig. 3.1 Top View

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Fig. 3.2 Front View

Fig. 3.3 Side View

Fig. 3.4 Isometric View

The sketches and the 3D model were created having no idea about the dimensions, or any type of geometric detail. These were drafted using a pure artistic approach. The next section explains the reasons for going for this unconventional layout.

3.2 Reasons for Triple Hull Layout

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A key motive for the decision of this configuration was differentiation, and especially for a layman’s eye. Other important reasons for the ruling are stated as under:

1. The theory of similarity predicts reduction of the payload capability when the dimensions of an aircraft are increased. Application of multi–fuselage aircrafts can improve the situation. (Ref. 2)

2. Separation of a large central mass into distributed outboard masses substantially alleviates wing bending loads (Ref. 3)

3. About 8% reduction in aircraft empty weight (Ref. 1)4. Better fuel efficiency; require smaller engines as that of a comparable single–

fuselage design (Ref. 3)5. Allows high aspect ratios to be used (Ref. 3)6. Disliking of twin–fuselage arrangement by the pilots (Ref. 1)

Having these augmented benefits, I opted for this layout and froze my concept.

Chapter 4

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AS DRAWN CONFIGURATION

4.1 Initial Weight Sizing

Initial Weight Sizing was carried out following the 3 rd Chapter of the textbook. The subsequent sections explain the procedure. Only the results are presented here. For complete steps of calculation, see Appendix A. In the initial weight sizing, it is assumed that fuel fraction for the descent segment equals unity.

4.1.1 Empty Weight Estimation

The average of the take off weights of the reference aircrafts were taken as the sized TO weight.

WTO)A320 = 162,000 lbf WTO)B737 = 174,200 lbf

Therefore, Sized WTO = 168,000 lbf

From Table 3.1: We/Wo = 0.496

8% empty weight saving due to triple fuselage layout results in:

We/Wo = 0.456

4.1.2 Fuel Weight

The weight fractions of the mission segments are presented as:

W1/Wo = 0.970W2/W1 = 0.985 W3/W2 = 0.834W4/W3 = 1W5/W4 = 0.993W7/W5 = 0.995W8/W7 = 0.985W9/W8 = 0.983W10/W9 = 1W11/W10 = 0.993W12/W11 = 0.995

So, the weight fraction of the whole mission equals:

W12/Wo = 0.753

With 6% reserves:

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Wfuel/Wo = 1.06(1 – W12/Wo) = 0.262

4.1.3 Payload

Wpassengers = 27750 lbfWBaggage = 9000 lbf

Therefore, WPayload = 36750 lbf

4.1.4 Crew

WCrew = 1850 lbf

4.1.5 Calculation of Total TO Weight

After determining the fuel weight and empty weight fractions, along with the weight of crew and payload, the total takeoff weight of the aircraft came out after iterating the following equation for We/Wo and Wo.

The results were: Wo = 139369.907 lbf, with We/Wo = 0.461

4.2 Airfoil and Geometry Selection

Following considerations were kept in mind while selecting the aerofoil:

1. High Cl,max 2. High Mcr to avoid transonic drag 3. Cl,design close to 0.493 (estimated design CL of the aircraft)4. Sufficient thickness to cater for the fuel weight

The aerofoil thus selected was NASA SC(2)–0610 (Mcr ≈ 0.81). No supercritical aerofoil was available for design Cl of 0.5 with the desired thickness. The supercritical aerofoil was selected, as it is effective in reducing wave drag, which would be there in the transonic regime.

4.2.1 Wing Geometry

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AR 9.412LE Sweep 30o

Incidence 1o

Twist 3o

Taper Ratio 0.7

Mid wing was selected with a dihedral of 1o.

4.2.2 Tail Geometry

T–Tail was selected for its added effectiveness and efficiency over the other configurations.

AR Taper Ratio

LE Sweep

Horizontal Tail 4.5 0.5 35o

Vertical Tail 1.2 0.9 40o

4.2.3 Tail Airfoil

Following considerations were taken into account while selecting the tail aerofoil:

1. Symmetric2. Higher Cl,max than wing3. Higher Mcr

than the wing4. 10% less thicker than the wing’s aerofoil

The airfoil thus selected was NACA SC(2)–0010. There was no thinner airfoil available in the category.

4.3 Design Point Selection

Design point means the selection of T/W and W/S for the aircraft. The design point for the aircraft was selected based on the study of 5th Chapter of the textbook. The T/W was selected by the Thrust Matching Technique, as

T/W = 0.278 (greater than that obtained from Table 5.3)

The wing loading was calculated for each mission requirement, and the detailed calculations are available in Appendix A. For the selected T/W, the wing loadings for the various requirements were obtained as:

Requirement W/S

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(lbf/ft2)Stall Speed 108.042

Takeoff Distance 179.254Landing Distance 106.242

Cruise 92.050Loiter 21.488

Absolute Ceiling 114.005

From Table 5.5, the value should be 120 lbf/ft2. The value obtained for loiter, the least important and also very low, was dropped. Also, that obtained for cruise was significantly less than 120 lbf/ft2. Keeping in view the 3rd Para of Article 5.4 (P 110), the selected W/S is the one determined by the absolute ceiling, i.e.

W/S = 114.005 lbf/ft2 (about 5% less than 120 lbf/ft2)

In T/W and W/S, T refers to the uninstalled, static and maximum thrust of the engine; W refers to the maximum takeoff weight of the aircraft; and S refers to the reference area of the wing.

4.4 Revised Weight Sizing

This section follows the calculation practices of the 6 th Chapter of the textbook. Here, the weight calculations were refined, and were more accurate than those determined in the initial sizing phase, as there was more information available.

4.4.1 Empty Weight Fraction

With guessed Wo = 168,000 lbf, We/Wo = 0.516 (from Table 6.1); and incorporating the benefit of empty weight reduction for the multi fuselage design gives:

We/Wo = 0.475

4.4.2 Fuel Weight

The fuel weight fractions for the mission segments were obtained as:

W1/Wo = 0.98W2/W1 = 0.979 W3/W2 = 0.849W4/W3 = 0.991W5/W4 = 0.993W7/W5 = 0.995W8/W7 = 0.986W9/W8 = 0.984W10/W9 = 0.995W11/W10 = 0.993

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W12/W11 = 0.995

As there is no payload drop and combat phase in the mission profile, fuel weight can be estimated as a fuel fraction.

W12/Wo = 0.761

With 5% reserves and 1% trapped fuel: WFuel/Wo = 0.253

4.4.3 Payload and Crew Weight

WPayload = 36750 lbfWCrew = 1850 lbf

4.4.4 Final Results

Eq. 6.3 reduces to Eq. 6.1 when WPayload drop = 0. Therefore, Eq. 6.1 is used to perform iterations to find WTO; and the result is:

Revised WTO = 143686.55 lbf, with We/Wo = 0.478

4.5 Geometric Sizing

This section deals with the dimensions of the aircraft. The numbers here are based on the revised aircraft weight and the design point. The details can be seen in Appendix A. The configuration contained two tails and three fuselages.

Fuselage Length per fuselage 68.961 ftMax. Diameter 9.195 ft

Wing

Sref 1260.353 ft2

Span 108.915 ftcroot 13.614 ftctip 9.53 ft

M.A.C. 11.692 ft

Horizontal Tail

Span 41.914 ftcroot 12.419 ftctip 6.21 ftS 390.389 ft2

Vertical TailSpan 20.27 ftcroot 17.78 ftctip 16.002 ftS 342.377 ft2

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4.6 Control Surface Sizing

The dimensions of the aircraft’s control surfaces were determined similar to those obtained above.

Aileron croot 2.45 ftctip 1.715 ft

Elevator croot 3.105 ftctip 1.552 ft

Rudder croot 5.69 ftctip 5.121 ft

4.7 Conclusion

With control surface sizing done, the geometric configuration of the aircraft was ready, and prepared for the next step of the conceptual design process: the Analysis, after the selection of a suitable power plant unit, i.e. engine. The engine specifications and the data are presented in the 7th Chapter. Now, the selection of the optimized aircraft is presented in the next chapter, which will be followed by the various analyses of that aircraft in the subsequent chapters of this report.

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Chapter 5OPTIMIZATION

5.1 Background

The analyses of the as drawn configuration were carried out after deciding the design point. The computations for those analyses were done in the Microsoft Excel Spreadsheets. After having the analyses (aerodynamic, propulsive and structural) done, the empty weight of the aircraft was found out using the more accurate methods of its evaluation, presented in the textbook (Eq. 15.25 to Eq. 15.45). The aircraft was then converged for the empty weight obtained, and the one got earlier in the revised weight sizing. The performance of this converged aircraft (now no more as–drawn) was then evaluated. The performance results show that whether the requirements/specifications have met or not.

All the parameters in the analyses kept on changing, the weights, areas, lift and drag coefficients, etc. except the design point of the aircraft. This design point is the prime parameter in the aircraft performance. A flowchart summarizing the scheme of work is shown as:

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Selection of Design Point

Analysis Aerodynamics Propulsion Structure

Converged Aircraft

Stability + Performance

Comparison of the required values with the achieved values

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This iteration was carried out for each aircraft configuration, by varying the design point. The design point was varied by +10% of the initial values of T/W and W/S. A sizing matrix was made to get an overview of the results.

5.2 Sizing Matrix

For preparing the sizing matrix, the design space was enlarged to 10% on both the sides of the original design point, to get eight more design points. A 3 X 3 matrix is thus obtained as:

T/W

W/S (lbf/ft2) 0.278 0.2502 0.306

114.005Ps @ 43k ft = -1.847

STO = 7603.95 SLanding = 6930.864WTO = 125619.6

Ps @ 43k ft = -9.114STO = 8353.814

SLanding = 7217.053WTO = 125210.1

Ps @ 43k ft = 5.461STO = 7163.32

SLanding = 6682.94WTO = 125996.8

102.6045Ps @ 43k ft = 3.503

STO = 6870.45SLanding = 6515.77WTO = 125964.6

Ps @ 43k ft = -3.77STO = 7535.45

SLanding = 6773.83WTO = 125551

Ps @ 43k ft = 10.80STO = 6479.66

SLanding = 6292.07WTO = 126345.6

125.406Ps @ 43k ft = -7.279

STO = 8338.15SLanding = 7342.16WTO = 125310.4

Ps @ 43k ft = -14.54STO = 9175.57

SLanding = 7656.38WTO = 124904.5

Ps @ 43k ft = 0.026STO = 7846.44

SLanding = 7069.95WTO = 125684.3

The Ps values are in ft/s, STO and SLanding in ft, and WTO in lbf.

RequirementsPs @ 43k ft = 0STO < 7000 ft

SLanding < 7000 ft

The WTO values are the converged ones, after the 15% usage of composites in the aircraft’s structure, thus reducing the empty weight of the aircraft by 23% than the comparables (8% due to multi hull layout).

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The computations of these results are available with the author in the form of MS Excel Spreadsheets. After making the sizing matrix, the constraints diagram was finalized to select the best possible aircraft, meeting the given requirements marginally.

5.3 Constraints Diagram and Selection of Optimized Aircraft

Three constraints of the absolute ceiling, TO distance and landing distance were included in the constraints diagram.

90 95 100 105 110 115 120 125 1300.24

0.25

0.26

0.27

0.28

0.29

0.3

0.31

Ps @ 43k ft, 0.7 MTO DistanceLanding Distance

W/S (lbf/ft2)

T/W

The available design space is the area of the graph enclosed within the three lines. The rest of the area is constrained by the lines. The final designed point chosen for the optimized aircraft was:

This was the point in the available design space where the T/W was maximum, and wing loading was minimum. At this design point, all the requirements have met, quite marginally.

5.4 Conclusion

With the optimization done, we now have the final aircraft at hand, which is the best possible configuration for the given specifications. The following chapters are based on this optimized aircraft. All the computations and procedures that have led to the convergence and optimization, are available as spreadsheets with the author.

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T/W = 0.278W/S = 108 lbf/ft2

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Chapter 6AERODYNAMICS

6.1 Introduction

This chapter deals with the aerodynamic analysis of the optimized aircraft. Here, the lift and drag estimates have been made.

6.2 Lift

The lifting characteristics of the aircraft were evaluated using the analytical methods given in the textbook. The results are presented here.

6.2.1 Lift curve slope versus Mach No.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

1

2

3

4

5

6

Mach No.

dCL

/ dα

(p

er ra

d)

The lift curve slope in the incompressible regime is equal to 3.707 per radian. This is also supported by the Fig. 12.5 of the textbook.

6.2.2 Maximum Lift Coefficient

The maximum lift coefficient of the aircraft in the clean configuration comes out to be 1.153. With the usage of High Lift Devices, it boosts up to 2.374. Triple slotted flaps and LE slats have been employed as the HLDs. The plot is shown on the next page.

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0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

0.5

1

1.5

2

2.5

Clean With HLDs

Mach No.

Max

CL

6.2.3 Stall Angle

In incompressible regime, the aircraft stalls at 18.168 degrees, while at the maximum lift coefficient, the stall angle decreases to about 16.7o. The variation of stall angle with Mach number is shown in the following graph.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

2

4

6

8

10

12

14

16

18

20

Clean With HLDs

Mach No.

Stal

l AoA

(deg

)

6.3 Drag

Both the parasite and induced drags were calculated for the aircraft. These estimates depend upon the wetted area of the aircraft, which was calculated after getting the exposed areas of the aircraft components from the CAD model and summing them up. The wetted area of the aircraft comes out to be 10265.23 ft2. At transonic speeds, the wave drag has also been calculated and included in the parasite drag coefficient.

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6.3.1 Parasite Drag Coefficient

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

0.005

0.01

0.015

0.02

0.025

0.03

0.035

SL15,000 ft35,000 ft43,000 ft

Mach No.

CDo

As shown in the above graph, the zero lift drag coefficient of the aircraft equals 0.0217 at sea level, Mach = 0.1.

6.3.2 Lift due to Drag Factor (K)

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

0.02

0.04

0.06

0.08

0.1

0.12

0.14

Uptill CL=0.51CL=0.98CL=1.44CL=1.9CL=Clmax=2.37

Mach No.

K

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6.3.3 0% and 100% K

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.90

0.05

0.1

0.15

0.2

0.25

0.3

Ko100% K

Mach No.

K

The increase in the 100% K value in the transonic regime can be seen.

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Chapter 7PROPULSIVE ANALYSIS

7.1 Introduction

From the design point information for the T/W and WTO, the engine of the aircraft should be capable of producing uninstalled thrust of 34971.38 lbf at sea level. As mentioned earlier in the morphological matrix, there are two engines. So, each engine should produce 17485.69 lbf.

7.2 Engine Selection

Rubber engine sizing was done. The engine selected is Rolls Royce Tay 650 (installed on Fokker 100).

The specifications of this engine are:

Thrust at sea level 15,100 lbfBypass ratio 3.07Inlet mass flow 422 lbm/sWeight 2949.22 lbfLength 7.9 ftFan diameter 3.75 ft

Based on the thrust required and the engine data available, the scaling factor of 1.158 was used to scale the rubber engine. The specifications of the rubber engine hence become:

Thrust at sea level 17,485.69 lbfBypass ratio 3.07Inlet mass flow 488.673 lbm/sWeight 3465.64 lbfLength 8.38 ftFan diameter 4.04 ft

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7.3 Engine Data

The data for the Rolls Royce Tay 650 engine was obtained from Ref. 1 (Chapter 9). The variation of thrust and TSFC with Mach number and altitude is given below.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

2000

4000

6000

8000

10000

12000

14000

16000

Thrust VS Mach No.

SL10,000 ft20,000 ft30,000 ft40,000 ft45,000 ft

Mach No.

Thru

st (l

bf)

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

0.2

0.4

0.6

0.8

1

1.2

TSFC VS Mach No.

SL10,000 ft20,000 ft30,000 ft40,000 ft45,000 ft

Mach No.

TSFC

(lbm

lbf-1

hr-

1)

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7.4 Rubber Engine Data

This was obtained after incorporating the scaling factor in the original engine’s data. The TSFC curves remain the same.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

2000400060008000

100001200014000160001800020000

Uninstalled Thrust VS Mach No.(Rubber Engine)

SL10,000 ft20,000 ft30,000 ft40,000 ft45,000 ft

Mach No.

Thru

st (l

bf)

7.5 Thrust Corrections

A 6% loss in the thrust resulted from the bleed losses. Catering this into the uninstalled engine data, the net propulsive force was obtained as shown below. This data was used further for the performance and stability calculations.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10

2000400060008000

1000012000140001600018000

Installed Engine Thrust VS Mach No.(Rubber Engine)

SL10,000 ft20,000 ft30,000 ft40,000 ft45,000 ft

Mach No.

Thru

st (l

bf)

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7.6 Capture Area Calculations

The capture area of the engine was calculated using the Fig. 10.16 of the textbook. With the engine mass flow known, and the design Mach number as 0.8, the Capture Area comes out to be 12.217 ft2.

From Eq. 10.16 and 10.17, the capture area comes out to be 11.281 ft2. This value is more accurate than that shown above.

7.7 Fuel System

The fuel used is JP–5. The total fuel volume comes out to be 4680.36 ft3. The fuel is stored in the wing and the fuselages.

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Chapter 8STRUCTURE AND COMPONENT WEIGHTS

8.1 Empty weight estimation from the component weights

Up till now, the empty weight was calculated from the empirical and statistical equations. This time, the empty weight was computed from the summation of the individual weights of the aircraft components, the equations for which have been provided in the 15 th

Chapter of the textbook. The following table shows the weights of the aircraft components. All readings are in lbf.

Wing 13670.852Horizontal Tails 3861.70448

Vertical Tails 2625.59761Fuselages 34742.4383

Main Landing Gear 1812.87298Nose Landing Gear 888.83839

Nacelles 1303.08612Engine Controls 65.1688

Pneumatic Starter 140.202864Fuel System 1404.94429

Flight Controls 1320.16771APU 2200

Instruments 417.992903Hydraulics 205.830709Electrical 686.142738Avionics 1689.4102

Furnishings 1237.81227Air Conditioning 2467.28224

Anti Ice 251.59268Handling Gear 37.738902

Total Empty Weight 71029.676278% Reduction due to Multi Fuselage

Layout 65347.3021715% Reduction due to the use of

Composites 55414.512

Hence, the empty weight of the aircraft comes out to be 55414.512 lbf. Furthermore, this is the weight of the aircraft when converged: The total weight of the aircraft is 125796.34 lbf, with We/Wo = 0.4405.

The multihull configuration along with the use of composite materials reduce the empty weight of the aircraft by 23%. This gives the aircraft a huge competitive advantage. This has resulted in the use of smaller engines, thus reduced drag and low fuel consumption. All the computation were performed in the MS Excel Spreadsheets.

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8.2 V–n Diagram

The plot between the load factor and equivalent air speed at sea level was made to see the structural limits of the aircraft. Cruise weight was considered for the calculations. The load factor limits were taken as:

Posiive nmax 3.5Negative nmax –1.5Positive Ultimate Load Factor 5.25Negative Ultimate Load Factor –2.25Factor of Safety 1.5

0 100 200 300 400 500 600 700

-3

-2

-1

0

1

2

3

4

5

6

V-n Diagram @ SL

V-n DiagramNegative Gust Load

EAS (ft/s)

Load

Fac

tor

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Chapter 9CREW STATION, PASSENGERS AND SPECIAL CONSIDERATIONS

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