CTA1 - GROUP BUSINESS DESIGN PROJECTPHASE 3 - AIRCRAFT PERFORMANCE

Baba KakkarAerospace Engineering

May 14th 2015

Executive Summary

Since world war II the demand for commercial aircraft has increased considerably, and inrecent years so has the aerospace technology. Therefore, civil aircraft design has become avery extensive process. The 2015 Airbus design project involved designing a medium-rangeairliner for the market segment between the existing short-range and long-range products,competing against the Airbus A330 and B767 re-engined. Blue sky aviation has come upwith an aircraft design called Azure, that meets the Airbus requirements. Azure has apassenger capacity of 274 in a two class layout for a design range of 5500nm, with a MTOWof 195 tons and maximum payload of 42.4 tons.

This report has analysed the performance characteristics of the aircraft while ensuring theoperations complied with the aerospace regulations. In order to carry out the required cal-culations, aerodynamic and propulsive characteristics were needed. To achieve this T0/WTO

and wing loading WTO/S was sized by constructing a constraints diagram against; time toclimb to ICA, cruise at ICA, approach at 145kts or less and take-off at second segmentclimb. A T0/WTO and wing loading WTO/S of 0.316 and 6244 (N/m2) was chosen, whichwould improve the aircraft’s fuel efficiency. With a selected wing area and engine size, theaircraft performance were assessed.

The flight profile provided by Airbus was used to work through the mission phase andreserves, to ensure the aircraft operations met the Airbus and CS25 requirements. Theclimb profile was optimised by climbing at the maximum excess power speed at differentaltitudes. Time to climb, distance travelled and fuel burnt to reach ICA were; 22.7 mins,150nm and 2.72 tons respectively. The initial cruise altitude was decided by optimising themach number and the ability of the aircraft to climb to 37,000ft. It was found that theoptimum mach number was 0.86 but this required the aircraft to climb between 37,000ftand 38,000ft. This was not possible as firstly at end of climb the aircraft would reach itsservice ceiling and also it was not able to climb to that altitude. Therefore cruise machnumber of 0.85 and ICA of 35,000ft were chosen. The cruise profile was then optimised byconsidering the cruise climb mode and assessing the best weight to change flight to keepa maximised (L/D). Cruise time, range and fuel burnt achieved were; 10.7 hours, 5215nmand 44.4 tons at a cruise climb mode respectively. The descent performance was not limitedby engine performance, therefore the fuel economy was not a critical factor as the fuelconsumption was low. The descent fuel, range and time achieved were; 112kg, 135nm and23 mins respectively.

Take-off and landing performance was also assessed to ensure they met the Airbus re-quirements. These distances were analysed at various WAT and runway conditions, whichprovided a better understanding of the aircraft and would also be helpful for airline compa-nies. At MTOW and ISA SL a take-off field length of 2550m was achieved, which is similarthe A330-300 even with a higher MTOW. Airbus requirement of TOFL was also met atISA +15◦. The landing field length was analysed at MLW and ISA SL, which was foundto be 1830m. However, this distance could be reduced by using thrust reversers or dragparachutes. A balanced field length analysis was assesed to consider the two possibilitiesif an engine failed during take-off; accelerated continue distance and accelerated stop dis-tance. This determined the decision speed V1 of 132kts, BFL of 2,400m and other significantspeeds.

A payload-range diagram was analysed to show the relationship between between range,payload and fuel. This determined the max payload, max economic and ferry range as5400nm, 7117nm and 7670nm respectively. In comparison to the competitor aircrafts, Azurecould fly at a longer economical range. In addition, also concluded from the flight profilewas the relationship between the block fuel at different range and payloads. This was usedas an input to assess how well Azure performed against its competitors at a mission rangeof 3000nm. The block fuel and time at 3000nm was found to be 28.1 tons and 6.5 hours. Itwas concluded that the Azure performed 5% better than its competitor A330 neo and 11%better than the B767 re-engined, against DOC.

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General Assembly Drawing

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Aircraft Layout

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Cabin Layout

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Technical Specification

Operational item’s

AccommodationMax seat (Single class) 373Two class seating 274No. Abreast 1st; 2nd;High 6;8;9Hold Volume (m3) 163Mass (Weight) (kg)Max. take-off (tons) 195Max. landing (tons) 170Zero-fuel (tons) 142Max. Payload (tons) 42.4Max. fuel payload (tons) 137Design payload (tons) 42.4Design fuel load (tons) 53.7Operational empty (tons) 98.2Weight RatiosMax. payload/Max. T/O 0.22Max. fuel/Max. T/O 0.28Max. landing/Max. T/O 0.87Fuel (litres)Block (tons) 49Reserve (tons) 5.6

Dimensions

FuselageLength (m) 61.3Height (m) 6.1Width (m) 6.1Finess ratio 2.85WingArea (m2) 306Span (m) 58MAC 5.91Aspect ratio 11Taper ratio 0.22Average (t/c) % 0.091/4 Chord Sweep 30.7High Lift DevicesTrailing edge flaps type Single slottedFlap span/wing span 0.595Area (m2) 33.6Leading edge flaps type slatsArea (m2) 38.6Vertical Tail

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Area (m2) 47.6Height (m) 8.86Aspect ratio 1.65Taper Ratio 0.351/4 Chord sweep 45Tail arm (m) 29.4Horizontal TailArea (m2) 80.6Height (m) 20.7Aspect ratio 5Taper Ratio 0.261/4 Chord sweep 30Tail arm (m) 28.7UndercarriageTrack (m) 10Wheelbase (m) 25.3Turning radius (m) 41.27No. of wheels (nose;main) 2;8Main wheel diameter (m) 1.05Main wheel width (m) 0.53NacelleLength (m) 6.38Max. width (m) 3.58Spanwise location 9.398

Performance

LandingMax. wing load (kg/m2) 6244Thrust/weight ratio 0.316Take-off (m)ISA sea level 2550ISA + 15C SL 2681Landing (m)ISA sea level 1830ISA + 15C SL 1894Speeds (kts/Mach)Vapp 145CLmax (T/O) 2.1CLmax (L/D @ MLM) 2.7Max cruiseSpeed (kts) 489Altitude (ft) 35,000Fuel consumption (kg/h) 0.0482Range (nm)Max. payload 5396Design range 5500Max fuel (+ payload) 7117Ferry range 7673

Contents

1 Introduction 1

2 Drag polar estimation 2

3 Aircraft sizing 23.1 Constraints diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

4 Take-off performance 4

5 Flight profile 75.1 Climb performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.2 Cruise performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115.3 Descent performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175.4 Block performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195.5 Flight profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

6 Landing performance 22

7 Balanced field length 25

8 Payload - Range 26

9 Specification comparison 27

10 Performance regulations 28

11 Technical drawing 30

12 Conclusion 30

13 Bibliography 31

A Drag polar estimations 33

B Constraints diagram 34

C Take-off performance 36

D Climb Profile 40

E Cruise Profile 44

F Descent Profile 48

G Flight profile 50

H Landing distance 54

I Balanced field length 55

J Payload - range 57

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K Performance regulations CS 25 58

L Performance regulations 58

M Technical drawing 65

N Aircraft layout 65

O Cabin layout 66

P General assembly 67

Notation

a Speed of sound m/s; acceleration, m/s2

CD Coefficient of dragCL Coefficient of liftCLG

Coefficient of lift in ground at zero angle of attackD Drag, NF General force, NFf Friction force, Ng Gravitational acceleration, m/s2

htr Transition height, mK Lift dependant drag factorL Lift, Nm Mass, kgM Mach numbern Load factorp pressure, N/m2

P Power, kWq Dynamic pressure, N/m2

R Radius, mS Gross wing area, m2

SG Ground distance mSTR Transition distance mSCL Clearance height mTSL Sea level static thrustt Time, s, min, hru Relative velocity, V/V mdR/C Rate of climb, m/sV Velocity, ktsVAPP Approach speed, ktsVEF Engine failure recognition speed, ktsVLOF Lift-off speed, ktsVmd Minimum drag speed, ktsVmp Minimum power speed, ktsVR Rotational speed, ktsVREF Reference landing speed, ktsV1 Decision speed, ktsV2 Take-off safety speed, ktsW Weight, NWTO Take-off weight, N

Greek symbols

α Angle of attack, deg, rad; thrust lapse rateβ Weight fractionγ Climb gradientµ Runway rolling coefficient of frctionθ Clearance angle, deg, radρ Air density, kg/m3

ω Fuel ratio, W1/W2

Abbreviations

ACD Accelerate continue distanceASD Accelerate stop distanceBFL Balanced field legnthCAA Civil airworthiness authorityCAS Calibrated airspeedCG Centre of gravityEAS Equivalent airspeedICA International cruise altitudetISA International standard atmosphereMLW Maximum landing weightMTOW Maximum take-off weightMZFW Maximum zero fuel weightOEW Operating weight emptySAR Specific air rangeTAS True airspeedTOFL Take-off field lengthTOW Take-off weight

List of Figures

1.1 Blue sky aviation group breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . 13.1 Constraints diagram showing design point (solid circle) . . . . . . . . . . . . . . . 34.1 Take-off distance breakdown at ISA + 15◦ . . . . . . . . . . . . . . . . . . . . . . 54.2 Change in take-off field length with different masses . . . . . . . . . . . . . . . . 75.1 Climb performance showing altitude against time, fuel and distance . . . . . . . 85.2 Maximum excess power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95.3 Maximum rate of climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105.4 TAS and EAS profile during climb . . . . . . . . . . . . . . . . . . . . . . . . . . 115.5 Cruise mode profiles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125.6 L/D against Cl (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135.7 ML/D optimisation (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . . . 145.8 Minimum power cruise climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

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5.9 Optimisation of flight level during cruise . . . . . . . . . . . . . . . . . . . . . . . 165.10 Carpet plot at cruise (Fletcher, 2015) . . . . . . . . . . . . . . . . . . . . . . . . 175.11 Range function at constant altitude cruise . . . . . . . . . . . . . . . . . . . . . . 175.12 Descent performance showing altitude against time, fuel and distance . . . . . . 185.13 TAS and EAS profile during descent . . . . . . . . . . . . . . . . . . . . . . . . . 195.14 Block fuel and range at different payloads . . . . . . . . . . . . . . . . . . . . . . 205.15 Direct operating cost comparison with competitors (Hassan, 2015) . . . . . . . . 205.16 Optimum flight profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226.1 Breakdown of the landing distance . . . . . . . . . . . . . . . . . . . . . . . . . . 237.1 Balanced field length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 258.1 Payload - Range diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268.2 Mass breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27A.1 Drag polar during cruise (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . 33A.2 Drag polar during take-off (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 33A.3 Drag polar during landing (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 34A.4 Drag polar during braking (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 34C.1 Change in CDgear

with flap deflection (Nicolai, 2011) . . . . . . . . . . . . . . . . 37C.2 Change in CDflap

with type and deflection of flap Nicolai, 2011 . . . . . . . . . . 37M.1 Leading edge slat drawing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65M.2 Leading edge slat track mechanisim . . . . . . . . . . . . . . . . . . . . . . . . . . 65

List of Tables

2.1 Estimation of drag polar during stages of flight (Dawson, 2015) . . . . . . . . . . 23.1 Wing loading and thrust to weight ratio comparison . . . . . . . . . . . . . . . . 34.1 Comparison of take-off distances . . . . . . . . . . . . . . . . . . . . . . . . . . . 44.2 Ground run results breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64.3 Take-off speeds breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65.1 Summary of results during climb . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.2 Phase 1 climb results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85.3 Phase 2 climb results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105.4 Design range breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115.5 Cruise mode comparison against cruise climb . . . . . . . . . . . . . . . . . . . . 125.6 ML/D against mach number . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.7 Summary of results during cruise climb . . . . . . . . . . . . . . . . . . . . . . . 145.8 Summary of results during stepped cruise . . . . . . . . . . . . . . . . . . . . . . 155.9 Summary of results during descent . . . . . . . . . . . . . . . . . . . . . . . . . . 185.10 Summary of results for reserve phase . . . . . . . . . . . . . . . . . . . . . . . . . 215.11 Fuel and time breakdown of flight profile . . . . . . . . . . . . . . . . . . . . . . . 216.1 Comparison of landing distances . . . . . . . . . . . . . . . . . . . . . . . . . . . 226.2 Summary of Airborne distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . 236.3 Summary of free roll distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 246.4 Summary of braking distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 247.1 Summary of balanced field length . . . . . . . . . . . . . . . . . . . . . . . . . . 258.1 Summary of payload-range diagram . . . . . . . . . . . . . . . . . . . . . . . . . 269.1 Airbus design specification comparison . . . . . . . . . . . . . . . . . . . . . . . . 2810.1 take-off distances at different surface runways . . . . . . . . . . . . . . . . . . . . 2810.2 OEI climb requirement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

B.1 Constraints analysis detailed results . . . . . . . . . . . . . . . . . . . . . . . . . 36C.1 Friction coefficient for various surfaces (Nicolai, 2011) . . . . . . . . . . . . . . . 38C.2 Ground run during take-off, detailed results . . . . . . . . . . . . . . . . . . . . . 38C.3 Take-off distance, detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . . 40D.1 Take-off, initial climb and acceleration detailed results . . . . . . . . . . . . . . . 41D.2 Phase 1 climb detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42D.3 Phase 2 climb detailed reults . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42D.4 Rate of climb at different altitudes . . . . . . . . . . . . . . . . . . . . . . . . . . 43E.1 Acceleration to cruise speed results . . . . . . . . . . . . . . . . . . . . . . . . . . 44E.2 Cruise climb detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45E.3 Stepped cruise 2000ft, constant altitude detailed results . . . . . . . . . . . . . . 46E.4 Stepped cruise 2000ft, climb detailed results . . . . . . . . . . . . . . . . . . . . . 46E.5 Stepped climb 1000ft, constant altitude cruise detailed results . . . . . . . . . . . 47E.6 Stepped cruise 1000ft, climb detailed results . . . . . . . . . . . . . . . . . . . . . 47E.7 Constant altitude cruise at 35,000ft detailed results . . . . . . . . . . . . . . . . . 48E.8 Constant altitude cruise at 37,000ft detailed results . . . . . . . . . . . . . . . . . 48F.1 Descent to 10,000ft detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . 49F.2 Descent to 1,500ft detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . 50G.1 Climb to 20,000ft during reserves . . . . . . . . . . . . . . . . . . . . . . . . . . . 51G.2 Constant altitude cruise at minimum drag velocity . . . . . . . . . . . . . . . . . 51G.3 Descent profile to 1,500ft during reserves . . . . . . . . . . . . . . . . . . . . . . . 52G.4 Start up and taxi out results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52G.5 Approach and landing results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53G.6 Taxi in results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53G.7 En-route allowance results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53G.8 Overshoot results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53H.1 Approach distance detailed results . . . . . . . . . . . . . . . . . . . . . . . . . . 54H.2 Ground braking distance detailed results . . . . . . . . . . . . . . . . . . . . . . . 55I.1 ASD detailed results part 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56I.2 ASD detailed results part 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56I.3 ACD detailed results part 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56I.4 ACD detailed results part 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57J.1 Max payload, economy and ferry range detailed results . . . . . . . . . . . . . . . 57

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1 Introduction

The 2015 Airbus design project involved designing a medium-range airliner for the market seg-ment between the existing short-range and long range single-aisle products, with a passengercapacity between 200-300, in a 2 class layout. The main key driver was a 15% reduction of directoperating cost (DOC) compared to the current 2010 state of the art, B767 re-engined and A330Neo, with a entry into service at 2025. The Blue sky aviation (CTA1) group was made up of 10members responsible for a technical and business role.

Figure 1.1: Blue sky aviation group breakdown

This report will highlight the final performance characteristics of the aircraft designed by Blue skyaviation (CTA1) named ’Azure’ and its economic performance as a transport vehicle, to meeta design range of 5500nm. To asses the performance of the aircraft, calculation of quantitieswere assessed, such as; rate of climb, maximum speed, distance travelled, mass of fuel andlength of runway required for take-off and landing (Mair, 1992). The safety parameters affectingthe performance was also taken into consideration to ensure safe operation handling of theaircraft.

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2 Drag polar estimation

The drag polar was estimated by the aerodynamicist and were interpolated at three stages; take-off, landing and cruise. These were then used to simplify the performance calculations. The dragcount for each component were cross checked with current aircrafts to see if they were reasonable.Table 2.1 shows the drag polar for the different stages and the corresponding graph in appendixA.

Take-off Cruise Landing Ground run with spolierCD0 0.0141 0.0155 0.0752 0.1644K 0.0483 0.0413 0.0460 0.0202

Table 2.1: Estimation of drag polar during stages of flight (Dawson, 2015)

3 Aircraft sizing

3.1 Constraints diagram

A constraints diagram was constructed to asses the technical constraints and the influence tothe design, shown in figure 3.1. This was then used to determine the final thrust-to-weight ratioand the wing loading. The four performance constraints that the aircraft was designed againstwere:

1. Approach speed during landing of 145kts or less2. Take-off from a 3000m field length3. Climb to ICA with 1.1% climb gradient4. Cruise at 0.85 mach at ICA5. Take-off with OEI

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Figure 3.1: Constraints diagram showing design point (solid circle)

Azure B767-300 B787 A330-300T0/W0 0.316 0.307 0.273 0.282W0/S (N/m2) 6244 6283 7637 5863S m2 306

Table 3.1: Wing loading and thrust to weight ratio comparison

The design point was chosen to minimise T0/W0 and the W0/S to give a reduced DOC andincreased aircraft efficiency, so the engines were not over sized and the largest possible wing areawas chosen. The design point was chosen to give a small amount of clearway. The results canbe seen in table 3.1, with comparison to other competitor aircrafts. The method and full resultscan be found in appendix B.

I. Take-off two engine and OEI

At take-off it was assumed β = 1, to simplify the calculation. Due to increase velocity, the lossesin thrust were considered and α was found to be 0.8. This was plotted against a range of wingloadings, which were plotted on the constraints diagram. The aircraft was designed to use singleslotted flaps, which were utilised to give a maximum CL of 2.1.

To asses the take-off constraint for one engine out, the second segment climb was considered,which required climb gradient of 2.4%. The equation was formulated to not only take intoaccount the drag from the landing gear and flap, but also the windmill drag. OEI climb andtake-off will be discussed in more detail in section 10.

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II. Cruise and climb condition

After taking into consideration hybrid laminar flow over the nacelles and tail plane, a better(L/D) was achieved and it was decided, a CD0 of 0.0155 and K of 0.0413 was achievable. Atthe top of climb the throttle setting was set to RC40 and at cruise the pilot throttled back toRC35. The thrust lapse rate α at top of climb was 0.181, which dropped to 0.170 at startingcruise. The fuel fraction was found from the mission profile for both conditions which includedthe taxi in, take-off and climb of 98.6% of the MTOW.

4 Take-off performance

A conventional take-off performance was assessed in two phases; the ground run distance andthe airborne distance. Conventional aircrafts usually require considerably longer runways whichrestrict the number of airports it can access. Therefore, with shorter distances the aircraft willbe more effective economically and operationally. The landing performance will be discussed insection 6 and the requirements for engine failure will be discussed in section 7.

Azure B767-300 B787 A330-300Take-off ISA SL (m) 2550 2545 N/A 2320Take-off ISA+20◦ SL (m) 2725 2850 2678 2680Take-off ISA+15◦ SL (m) 2681

Table 4.1: Comparison of take-off distances

The results calculated show good similarities compared to the competitor aircrafts. A breakdownof the take-off distances are shown in figure 4.1 at ISA + 15◦ SL. The methods and detailed resultsare shown in appendix C.

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Figure 4.1: Take-off distance breakdown at ISA + 15◦

I. Ground distance

The biggest portion of the take-off distance contributed to the ground run, which was a total of1518m and took 34 seconds. To calculate the ground distance the model also included the dragfrom the landing gear, the flap deflection and the brakes-off friction force. The lift coefficientat ground run was found from the CL vs α graph, which was 0.731. In addition, the flap andlanding gear drag coefficient were cross checked with the graphs shown in appendix C. TheCDgear predicted by the aerodynamicist was found to be 0.004, however I believe this was verysmall compared to a medium/large transporter aircraft. During calculations this change wasfound to only decrease the take-off distance by 2%. It was assumed that the flaps were deployedbefore the ground run and the CDgear

was constant throughout the ground run. The loss inthrust and SFC was due to the increased velocity and this was taken into consideration whenintegrating between start of ground run to the rotation velocity. The rolling friction coefficientwas found from table D.4, as 0.015.

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AzureWeight (kg) 195 tonnesTE flap deflection 15 degRolling friction coefficient 0.015Ground run CL, CLG

(m) 0.73 (at α = 0 and flap deflection of 15 deg)Thrust with 2 engines (N) 604257 kN (At start of runway)δCDflap

0.015δCDgear

0.004Air density ρ (kg/m3) 1.225 kg/m2

Table 4.2: Ground run results breakdown

II. Rotation distance

The rotation distance contributed to 6% of the ground run distance, which was a total of 164mand took 2 seconds. The aircraft rotated on the ground until an angle of attack of 10◦ was reachedwhich gave a CL of 0.8CLMAX

. According to the GA, a rotation angle of 12◦ was available, hencethere would be no tail strike.

III. Transition distance

The transition distance was the second biggest phase of the take-off distance, which contributedto 673m and took 8 seconds. During the transition phase the aircraft flew at a constant velocityarc at radius R. The load factor was found to be 1.152, which gave a radius of 4934m. Howeverat this radius and takeoff velocity the clearance distance of 35ft was reached and exceeded, hencethe the climb distance was zero.

The total distance was then found to be the sum of the rotation, transition and ground distance.This summed up to be 2332m, but was multiplied by 115% to account for safety regulations.Performance regulations will be discussed in detail in section 10. If the effects of headwindand tailwind were considered, then it can be seen from the equation in section C, that with aheadwind of 5%, the ground speed would only have needed to be accelerated to VT0 − VW . Thismeant the ground run would have decreased by 10%, and if the tail wind were considered thenthe ground run would have increase by 10%. The new total take-off distance would then be2654m, which would still lie within the Airbus specification. Table 4.3 show a breakdown of thespeeds during the take-off analysis.

AzureStall speed VS (kts) 139Rotation speed VR (kts) 161Lift off speed VLO (kts) 167(L/D) when airborne 12.5

Table 4.3: Take-off speeds breakdown

In normal operations it was most likely that airline companies would not be a taking-off withthe maximum take-off weight. Therefore, figure 4.2 shows how the take-off field length varies atdifferent masses.

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Figure 4.2: Change in take-off field length with different masses

5 Flight profile

5.1 Climb performance

The climb profile was split into two phases; climb from 1,500ft to 10,000ft at constant 250ktsCAS, which was set by regulations and climb from 10,000ft to 35,000ft at 300kts CAS set byAirbus specification. However, the second phase of climb was allowed to be changed and wasoptimised so better climb rate could be achieved. The method and detailed results can be foundin appendix D.

AzureFuel burnt (kg) 2724Time to climb (min) 22.7Distance flown (nm) 150MTOW/ICA (top of climb) 0.986

Table 5.1: Summary of results during climb

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Figure 5.1: Climb performance showing altitude against time, fuel and distance

I. Climb - phase 1

The first climb phase contributed from 1,500ft to 10,000ft and acceleration to the design climbspeed. During this climb phase no optimisation could be possible as the climb speed was re-stricted. However, due a better (L/D) from the hybrid laminar flow, the rate of climb improvedwhen compared to phase 2B. At the end 10,000ft a (L/D) of 20 was achieved. Table 5.3 highlightsthe climb results during this phase.

AzureFuel burnt (kg) 664Time to climb (min) 4.5Distance flown (nm) 22Rate of climb at 10,000ft (ft/min) 2500(L/D) at top of climb 20.2Mach at top of climb 0.46T0/WT0 0.135MTOW/(top of climb) 0.997

Table 5.2: Phase 1 climb results

II. Climb - phase 2

The second climb phase contributed from 10,000ft to ICA of 35,000ft. Initially the aircraft wasdesigned to climb at constant 300kts CAS, however after plotting a graph of velocity against

9

power and (R/C), better results were achieved. The CAS was assumed to be equivalent to EAS,as the scale-altitude correction was smaller during climb. Figures 5.2 and 5.3 shows an exampleof the results after optimisation at 10,000ft, which were completed for each altitude at 5000ftstep climbs.

Figure 5.2: Maximum excess power

As illustrated on the graph, the best (R/C) were achieved when the there was maximum excesspower between the power available from the thrust and the power required from the drag. Thegraph also shows the maximum velocity the aircraft could climb at before more power wasrequired then available. This was then completed for each altitude at 5,000ft climbs until ICA,results for this phase can be seen in table 5.3. In addition, the mach at top of climb was alsoclear from the critical mach number of 0.88. As the thrust varied with airspeed, it was foundthat the minimum drag speed did not provide the optimum climb. Figure 5.4 shows how theTAS and EAS varied during the climb until ICA.

10

Figure 5.3: Maximum rate of climb

AzureFuel burnt (kg) 1936Time to climb (min) 17.6Distance flown (nm) 126Rate of climb at 35,000ft (ft/min) 375(L/D) at top of climb 20.2Mach at top of climb 0.76T0/WT0 0.058MTOW/(top of climb) 0.986

Table 5.3: Phase 2 climb results

11

Figure 5.4: TAS and EAS profile during climb

To achieve a lower DOC, block time was needed to be reduced which reflected slightly on thetime to climb (Jenkinson, 1999). At higher altitudes the aircraft would not have enough airdensity to produce higher thrust, hence the rate of climb would be reduced. This then reflectedon the time to climb and the distance flown, as seen by the shape of the graph in figure 5.2.However, the fuel burnt was kept constant throughout the climb by optimisation, to make sureexcess fuel was not burnt.

5.2 Cruise performance

To analyse the cruise performance, the cruise range was first identified for the design mission of5500nm. Using the calculated climb and descent range, the cruise range was found to be 5215nmas shown in table 5.4. It was decided to cruise at ICA of 35,000ft and the reasons for the choicewill be discussed below.

AzureDesign range (nm) 5500Climb range (nm) 150Descent range (nm) 135Cruise range (nm) 5215

Table 5.4: Design range breakdown

12

The aircraft was capable of cruising in three possible cruise modes, depending on air trafficcontrol; cruise climb, stepped cruise and cruise at constant altitude. These three modes wereanalysed, shown in figure 5.5 and summary of results shown in table 5.5. The methods anddetailed results are shown in appendix E.

Increase fromoptimum

Cruise climb fuel (tons) 44.43Cruise climb time (hr) 10.70Step cruise (2000ft) fuel (tons) 45.19 1.73%Step cruise (2000ft) time (hr) 10.84 1.32%Step cruise (1000ft) fuel (tons) 44.86 0.98%Step cruise (1000ft) time (hr) 10.77 0.72%Constant altitude (35,000ft) fuel (tons) 45.79 3.07%Constant altitude (35,000ft) time (hr) 10.70 0.00%Constant altitude (37,000ft) fuel (tons) 44.49 0.14%Constant altitude (37,000ft) time (hr) 10.70 0.00%

Table 5.5: Cruise mode comparison against cruise climb

Figure 5.5: Cruise mode profiles

13

I. Mach optimisation

To optimise the mach during cruise at ICA, a graph of ML/D was plotted against CL and machnumber, shown in figure 5.7. It can be seen that the optimum mach number was between 0.86and 0.87, however this required a CL between 0.55 and 0.60. This would mean the aircraft wouldneed to climb between 37,000ft and 38,0000ft, and at the end of cruise the altitude would behigher then the service ceiling. Therefore the optimum mach number at initial cruise altitude of35,000ft would be at 0.85. In addition, the aircraft would not be able to produce a positive climbgradient at 37,000ft, as it was designed on the constraints diagram. This would then require ahigher engine size, which would increase the MTOW. Table 5.6 shows a comparison of cruisemach numbers with competitor aircrafts.

Figure 5.6: L/D against Cl (Dawson, 2015)

14

Figure 5.7: ML/D optimisation (Dawson, 2015)

Azure B767-300 B787 A330-300Cruise speed (kts) 490 489 488 489ICA (ft) 35000 39000 37000 33000Cruise Mach 0.85 0.85 0.85 0.86

Table 5.6: ML/D against mach number

II. Cruise climb

During cruise climb the air pressure decreases with increasing altitude, which allows for thedecrease in aircraft weight as fuel was being burnt. The angle of attack was also kept constantthroughout the cruise, therefore (L/D) was constant at 19.5. The most optimum range duringcruise climb occurred at minimum power velocity, Vmp = 1.136Vmd. However flying at minimumpower velocity was not possible was shown in figure 5.8. It can be seen that the designed cruisespeed was only 70kts below the optimum. The results calculated during the cruise climb wasintegrated at 1000ft steps to allow for the loss in thrust and sfc shown in table 5.7.

AzureICA (ft) 35,000Final cruise altitude (ft) 40,000L/D 19.5VTAS (kts) 490VMP (kts) 560VMD (kts) 426

Table 5.7: Summary of results during cruise climb

15

Figure 5.8: Minimum power cruise climb

III. Stepped cruise

As cruise climb was the optimum climb condition for best range and fuel efficiency, the steppedcruise mode was iterated to behave in a similar way, as seen in figure 5.5. This was achievedby setting the range to match equal areas above and below the cruise climb profile. This thenallowed W/ρ to be kept close to the required value. The 2000ft step cruise only used 1.73% morefuel than the cruise climb, but the time was also increased. This was because when climbing at2000ft steps, the power available was reduced due to the altitude, and therefore time taken toclimb was longer. During climb the throttle setting was increased to RC40 and due to decreasein potential energy there was a slight loss in airspeed. The mach number during climb wasalso restricted to 0.84, which was assumed to be achieved after throttle change and increase inaltitude. In addition, the 1000ft step climb performed better, as W/ρ was kept closer to requiredvalue. Table 5.8 shows the summary of results.

AzureICA (ft) 35,000Final cruise altitude (ft) 39,000L/D at 39,000ft 19.8R/C at 39,000ft (ft/min) 348Climb Mach 0.84

Table 5.8: Summary of results during stepped cruise

A graph of VL/D was plotted against CL at constant mach of 0.85 for different altitudes toasses the best weight to change flight (shown in figure 5.9). At a CL of higher than 0.62, thecompressibility drag increases substantially and therefore the lift to drag would be become very

16

low. However, the CL would never reach that value as the cruise range would be complete andtherefore there would be no need to change to a flight level of 40,000ft.

Figure 5.9: Optimisation of flight level during cruise

IV. Constant altitude cruise

During this cruise mode the angle of attack decreased as weight decreased and the mach numberwas kept constant. As the cruise speed was greater than the minimum drag speed, then thedecrease in lift coefficient would decrease the drag, meaning thrust will also be decreased. Thismeant more fuel would be required in comparison to the other cruise modes (shown in table 5.5).At 35,000ft this was seen to be correct, however at 37,000ft this cruise mode seemed to be bettercompared to the step cruise modes.

In figure 5.10 and 5.11, at 0.85 mach and 37,000ft it can be seen that the sfc was lower comparedto other altitudes. As thrust decreases during the cruise the sfc improves, getting closer to itslow peak. This will not be the case for cruising lower than 37,000ft or higher, as sfc starts toincrease. In addition, figure 5.11 shows the relative velocity (Vmd/V ) against the range for theconstant altitude climb at a fuel fraction of 1.3. Between the initial cruise velocity and the finalcruise cruise velocity, optimum range was achieved. As minimum drag speed reduces duringcruise with aircraft weight the relative velocity increases, reaching the optimum range peak forbest efficiency. Cruising at any higher or lower velocity would decrease the optimum range andincrease the fuel burnt.

17

Figure 5.10: Carpet plot at cruise (Fletcher, 2015)

Figure 5.11: Range function at constant altitude cruise

5.3 Descent performance

Similar to the climb profile, the descent profile was split into two main phases: descent from40,000ft to 10,000ft, where descent was made at a airspeed of 300kts CAS and from 10,000ft to

18

1,500ft where descent speed was restricted to 250kts CAS. The equations shown in appendix Fwere used to calculate the results shown in table 5.9 and figure 5.12, which were integrated at5,000ft intervals.

AzureFuel burnt (kg) 112Time to descent to 1,500ft (min) 23Distance flown (nm) 135

Table 5.9: Summary of results during descent

Figure 5.12: Descent performance showing altitude against time, fuel and distance

The descent performance was not limited by engine performance, as the throttle setting was setback to RC20, flight idle and the engines produce residual thrust. Therefore, during descent thefuel economy was not a critical factor as the fuel consumption was low. However, the rate ofchange of cabin pressure was to be kept at low value to avoid any problems with passengers onboard. It was agreed that the rate of change of cabin pressure would not exceed 300ft/min atsea level. With the cabin pressure at 41,000ft set to be equivalent to 6,000ft, the time to descentto 1,500ft should not have exceeded 15 mins. To avoid not increasing the rate of climb the pilotmay have chosen to use flaps or spoilers to increase the drag on the aircraft to a suitable value.However, the engines were required to be in full working conditions in order to keep generatorsand hydraulic pumps active, so throttle setting could have been chosen to be increased. It wasdecided to descend at 300kts CAS as specified by Airbus, but initially at constant mach of 0.84to avoid the critical mach number. Figure 5.13 shows how the TAS and EAS varied during thedescent to 1,500ft.

19

Figure 5.13: TAS and EAS profile during descent

5.4 Block performance

The block performance was a really useful summary for airline operators for a quick estimate onhow much fuel was required for a specific mission. This determined the economic viability of theaircraft and the operating cost shown in figure 5.14.

This information was also important for the DOC calculations of a mission range of 3000nm,which was one of the hard Airbus requirements. Together with the finance team, the DOC, COCand seat mile cost were calculated for Azure and its competitor aircrafts, shown in figure 5.15.The block fuel was found to be a big driver when improving the aircrafts economy.

20

Figure 5.14: Block fuel and range at different payloads

Figure 5.15: Direct operating cost comparison with competitors (Hassan, 2015)

21

5.5 Flight profile

I. Reserve phase

A similar approach was taken for the reserves phase of the flight profile, where the methodscan be found in appendix G. Table 5.10 shows the summary of the results during the reservephase.

AzureApproach and landing, fuel (kg) 83Approach and landing, time (min) 5Taxi in, fuel (kg) 158Taxi in, fuel (min) 7En-route allowance, fuel (kg) 1804Diversion, fuel (kg) 2254Diversion, time (min) 45Holding, fuel (kg) 1494Holding, time (min) 30Total reserve, fuel (kg) 5634Total reserve, time (min) 80

Table 5.10: Summary of results for reserve phase

II. Flight profile

A summary of the flight profile can be see in table 5.11 and an example of the an optimised flightprofile was shown in figure 5.16.

AzureMission range, fuel (tons) 47.9Mission range, time (hrs) 11.5Flight fuel (tons) 48.2Flight time (hrs) 12Block fuel (tons) 49Block time (hrs) 12.2

Table 5.11: Fuel and time breakdown of flight profile

22

Figure 5.16: Optimum flight profile

6 Landing performance

A conventional landing performance was assessed in two phases; the ground run distance and theairborne distance. The ground run distance was the sum of the rolling distance and the brakingdistance. The methods described in appendix H have been used to calculate the results shownin table 6.1. The aircraft mass at the maximum landing weight was used to give the longestlanding distance.

Azure B767-300 B787 A330-300Landing distance ISA SL (m) 1830 1740 1520 1600Landing distance ISA SL + 20◦ (m) 1894 1740 1520 1600

Table 6.1: Comparison of landing distances

The results calculated show good similaraties compared to the competitor aircrafts. A break-down of the take-off distances was shown in figure 6.1, which was a specified in the Airbusrequirement.

23

Figure 6.1: Breakdown of the landing distance

I. Airborne distance

The airborne distance was the horizontal distance, to clear a 50-ft obstacle, where the velocitywas 1.3VS . A stall speed of 103kts was achieved from a CL of 2.7, which was optimised bythe aerodynamicist. During calculations, it was assumed that the gradient of the approach wasaround 3 degrees, which was similar for commercial aircrafts. The airborne distance contributedto 21% of the landing distance, which was a total was of 233m and took 6 seconds. To calculatethe air distance the model included the drag from the landing gear and the flap deflection at 10degrees. The results during the airborne distance can be seen in figure 6.2.

AzureWeight 147 tonnesTE flap deflection 10 deg(L/D) 3.6CDflap

0.015CDgear

0.004CLmax 2.7VS kts 103V50 135Air density ρ 1.225 kg/m3

Table 6.2: Summary of Airborne distance

24

II. Free roll distance

The free roll distance was the distance travelled while the pilot reduced the throttle setting toflight idle, deployed the spoilers and applied the brakes, which contributed to 10% of landingdistance with a total of 155m. This allowed the rate of descent and airspeed to be reduced toa safe value. This distance was assumed to last 3 seconds at the touchdown velocity, which wasequal to 1.15VS . Results from the free roll distance are shown in table 6.4.

AzureFree roll distance 184TE flap deflection 10 degCDflap

0.015CDgear

0.004CLmax 2.7VS kts 103VTD 135Air density ρ 1.225 kg/m3

Table 6.3: Summary of free roll distance

III. Braking distance

The braking distance covered the biggest portion of the landing distance, which was a total of750m and lasted 22 seconds. The ground lift coefficient CLG

was found from the CL vs α graph,as 1.366 at the landing configuration.

AzureTE flap deflection 10 degCDflap

0.015CDgear

0.004CDspoiler

0.006µ brakes on 0.3CLG

1.366(L/D) 6.8VTD 135Air density ρ 1.225 kg/m3

Table 6.4: Summary of braking distance

To decrease the landing distance a larger retardation force was required, which could be ob-tained by increasing the drag or decreasing the stall velocity. This meant with an higher CLmax ,shorter landing distances could be achieved, which would be more effective economically andoperationally.

25

7 Balanced field length

The balance field length determined the decision speed V1, from which the aircraft could continueto accelerate and take-off or stop to a halt when an engine failed. This was calculated by selectingan engine failure speed VEF from which the pilot had 1 second to make his decision and then 2seconds to act. As one engine fails at VEF the thrust was reduced and the drag was increased dueto windmilling of one engine. Figure 7.1 and table 7.1 shows the summary of the results.

Figure 7.1: Balanced field length

AzureVEF (kts) 130V1 (kts) 132Vfirstaction (kts) 136VR (kts) 155VTO (kts) 162Balanced field length (m) 2400

Table 7.1: Summary of balanced field length

26

8 Payload - Range

As the mission profile focused on time and fuel on a specific range and take-off, the payload-rangediagram looked at the balance of range, payload and fuel. Figure 8.1 shows the payload-rangediagram for the aircraft. Main results achieved from this diagram can be seen in table 8.1. Thiswas achieved with working with the technical integrator and the methods described in appendixJ.

Figure 8.1: Payload - Range diagram

Azure A330-300 B767-300OWE (tons) 98.2MTOW (tons) 195MLW (tons) 170Max payload (tons) 39.2Max payload range (nm) 5400 3888 3221Max economic range (nm) 7117 7056 5354Ferry range (nm) 7670

Table 8.1: Summary of payload-range diagram

The max economic range could achieved when the wing used up all its capacity, which was foundby the fuel systems manager. An increase in range would require a decrease payload, which woulddecrease the drag. However, any range past the economic range will not be economical and DOCwill decrease. If a short mission was required to be completed and the aircraft was needed toland below or at its maximum landing weight then fuel must be dumped. With this diagram,

27

airlines companies could asses at what payload and range the mission would be economical. Afull breakdown of the aircraft mass can be seen in figure 8.2.

Figure 8.2: Mass breakdown

9 Specification comparison

Table 9.1 shows the comparison of the Airbus specification and the requirements and how wellAzure performed against it. All specifications were met.

28

AzurePassenger Capacity (2 class) - 200-300 274

Design Range nm 5500 5500

Design Cruise Speed Mach 0.82 - 0.86 0.85

Time to climb (1500ft to ICA at ISA) mins. ≤ 30 21.7

Initial Cruise Altitude ft 35000 35000

Maximum Cruise Altitude ft 41000 40000

Approach speed (MLW, SL, ISA) kts CAS ≤ 145 145

Take Off Field Length (MLW, SL,ISA+15)

m 3000 2681

Landing Field Length (MLW, SL, ISA) m 2500 1895

One Engine Inoperative Altitude ft Result 20000

VMO/MMOkts CAS /

Mach360kts /M=0.89

360kts /M=0.89

Equivalent Cabin Altitude (at 41000ft) ft 6000 6000Turn-Around Time mins. Result 52

Airport compatibility limits -ICAO Code

EICAO Code

EACN (Flexible B) - 60 60

DOC target $/seat-nm2010 Stateof the art

minus 15%15%

ETOPS capability (at EIS) mins. 180 293Expected Entry Into Service Year 2025 2025

Table 9.1: Airbus design specification comparison

10 Performance regulations

The regulations from the CS.25, which affect the performance of the aircraft were constantlyreferred to. Most regulations were met and are found in appendix L with results shown be-low.

I. CS 25.105 Take-off

(a) - The take-off speeds discussed in CS25. 107 were assessed against most weights and tem-peratures.

(b/c) - Take off distances are assessed in table below on different surfaces

AzureTake-off ISA SL (m) concrete 2550Take-off ISA SL (m) wet concrete 3118Take-off ISA SL (m) soft turf 3342

Table 10.1: take-off distances at different surface runways

29

II. CS 25.107 Take-off speeds

(a) - VEF = 130kts was less than V1 = 132kts

(b) - V2,min = 162kts was bigger than 1.13VSR = 153kts

(c) - V2 = 167kts was bigger than 1.13V2,mins = 162kts

(d) - Vmu = 162kts

(e) - VR = 136kts was bigger than V1 = 132kts

(f) - VLOF = 135kts

III. CS 25.109 Accelerate stop distance

(a) - ASD = 2400m and was bigger than the distance to VEF with AEO at 1550m. The ASDwas greater than the distance from standing to highest speed reached during rejected take-off of1800m.

(b) - ASD at wet runway = 2850m was greater then the ASD at dry runway of ASD 2400m

(c) - The wet runway braking coefficient of friction for a smooth wet runway was defined from atextbook - NOT MET

(d) - It was assumed the wet runway had no extra treatments done - N/A

(e) - The brake coefficient used was reliable and suitable for a similar size aircraft

(f) - Thrust reversers were not included in the calculations

(g) - The landing gear drag was present throughout the ASD

(i) - Flight test data not available - N/A

IV. CS 25.111 Take-off path

(a) - The take-off distance was calculated by using the specification as the aircraft accelerated onthe ground to VEF , at which point the critical engine was made inoperative and after reachingVEF, the airplane was accelerated to V2

(b) - During the acceleration to speed V2 the landing gear was not retracted at speed of 134ktswhich was below VR = 136kts

(c) - The gradient during take-off was positive at each point. At 400ft the gradient with OEIwas 1.8% which was higher than 1.2%

(d) - The take-off path was calculated at segments of time and altitude, where the parameterswere kept consistent. In ground and out of ground effect were taken into consideration

V. CS 25.113 Take-off distance and run

(a) - The take-off path was 2218m, which was greater than the take-off distance of 2550m

30

(b) - The take-off distance on a wet runway was 2880m which was greater than the take-offdistance of 2550m

(c) - During take-off calculations a clearway was not included.

VI. CS 25.115 Take-off flight path

(a) - The take-off path started from 35ft above the take of surface

(b) - More than 0.8% gradient was achieved during initial climb and en-route climb. At 1,500fta climb gradient of 20% was available.

(c) - Climb gradient was reduced during the climb path with 0.83% at 35,000ft.

VII. CS 25.121 OEI climb

(a) - During take-off at take-off power and with the undercarriage down at OEI, the climbgradient was positive.

(b) - At second climb with take-off power with the undercarriage retracted the required climbgradient was met.

(c) - At final take-off at max continuous power and undercarriage up, the required climb gradientof 5.8% was met.

Segment 1 Segment 2 Segment 3Climb gradient 8.0% 5.5% 5.8%

Required gradient Positive 2.4% 1.2%(L/D) 10 9 9.5(R/C) 1.3 4.6 5.0

Table 10.2: OEI climb requirement

11 Technical drawing

The technical drawing of the leading edge and literature review can be found in appendixM.

12 Conclusion

In conclusion this report illustrated the final performance targets for Blue sky aviations air-craft, Azure, using the provided aerodynamics and propulsive characteristics. Assumptions werebrought to a minimum but kept consistent throughout the performance analysis. The perfor-mance characteristics met all Airbus specification, as well as most CS regulations.

The selection of the engine and wing loading were minimised by the construction of the constraintsdiagram. This provided adequate information for the group members to perform calculations

31

and finalise the design. Reasonable optimisation during climb and cruise profiles determinedthe the most fuel efficient profile the aircraft was capable of. Due to air traffic control is it iscommon for aircrafts not to cruise in the most efficient path, hence other cruise modes were alsoanalysed. It was concluded that other models do incur penalties on fuel burn and time but notseverely.

The performance model also assessed a few WAT analysis, however the model could be improvedby considering an more in depth WAT analysis for each performance characteristic. Airlinecompanies do not perform missions with maximum take-off or full the mission range, hencehaving this analysis could be really useful. The performance model also considered the effects ofthe block performance against take-off at different masses or payload and different range missions.This determined how well Azure performed it had cruise to a different range. Maximum economyrange was also assessed against payload, which helps airline companies manage the payload andrange to maximise profit. On the contrary, after flight test data more accurate and reliableperformance characteristics would be estabilished. Using this performance model and inputfrom other group members, the Azure was able to perform 11% better then the B767 re-enginedand 9% than the A330 neo.

13 Bibliography

Anderson, J. (1999). Aircraft performance and design. Boston: WCB/McGraw-Hill.

Dawson, Martin (2015). Aircraft aerodynamics - Phase 3. CTA1 group business design project,University of bath

Fletcher, Thomas (2015). Aircraft propulsion - Phase 3. CTA1 group business design project,University of bath

Hassan, Muby (2015). Direct operating cost - Phase 3. CTA1 group business design project,University of bath

Jenkinson, L., Simpkin, P. and Rhodes, D. (1999). Civil jet aircraft design. Reston, VA: Amer-ican Insitute of Aeronautics and Astronautics.

Lee, Jia Juan (2015). Technical integrator - Phase 3. CTA1 group business design project, Uni-versity of bath

Mair, W. and Birdsall, D. (1992). Aircraft performance. Cambridge [England]: Cambridge Uni-versity Press.

Mok, Tim (2015). Aircraft fuel systems - Phase 3. CTA1 group business design project, Univer-sity of bath

32

Shelby, M. (2000). Aircraft performance, Theory and practice. Great Britain: Arnold, VA:American Insitute of Aeronautics and Astronautics.

Shevell, R. (1983). Fundamentals of flight. Englewood Cliffs, N.J.: Prentice-Hall.

Nicolai, Leland (2011). Fundamentals of Aircraft and Airship Design, Volume I - Aircraft De-sign, American Institute of Aeronautics and Astronautics.

33

Appendices

A Drag polar estimations

The four graphs shown below were used to estimate the drag polar by interpolating between thelift coefficients at the different stages of flight path.

Figure A.1: Drag polar during cruise (Dawson, 2015)

Figure A.2: Drag polar during take-off (Dawson, 2015)

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Figure A.3: Drag polar during landing (Dawson, 2015)

Figure A.4: Drag polar during braking (Dawson, 2015)

B Constraints diagram

I. Approach speed

With the approach speed set by Airbus as ≤ 145kts, the following equation was rearranged togive (MLW/S) (Shevell, 1983):

VAPP ≥ 1.3 ×

√2(MLW/S)

ρCL,Max,Land(B.1)

This was then converted back to maximum allowable wing loading at take-off, W0/S, to beincluded in the constraints diagram.

35

II. Take-Off Field length

The thrust to weight ratio required to take-off was calculated by the following master equa-tion:

T0W0

=1.44β2

αSGgρCLmax

(WTO

S

)(B.2)

This equation was altered to give the relevant form to equate the drag polar provided by theaerodynamicist.

III. Climb to initial cruise altitude

The thrust to weight ratio required to climb to the initial cruise altitude at the design machnumber was calculated by the following master equation:

T0W0

=β

α

[CD0

(β/q)(WT0/S)+ k

β

q(WT0/S) +

1

V

dh

dt

](B.3)

This equation was altered to give the relevant form to equate the drag polar provided by theaerodynamicist.

β and α are found using the following relationship:

T = αTSL (B.4)

W = βWTO (B.5)

IV. Cruise at ICA at M 0.85

The thrust to weight ratio required to cruise at mach 0.85 at the initial cruise altitude wascalculated by the following master equation:

T0W0

=β

α

[CD0

(β/q)(WT0/S)+ k

β

q(WT0/S)

](B.6)

This equation was altered to give the relevant form to equate the drag polar provided by theaerodynamicist.

V. OEI Climb Constraint

To meet the airworthiness requirements of minimum climb gradients, take-off and climb to sec-ond segment with one engine out was added to the constraints diagram. This would determine

36

the thrust required to achieve the minimum climb gradient set by CS 25. To take into consid-eration loss of available thrust with one engine out, the following equation was used, (Shevell,1983).

T0W0

=β

α

(n

n− 1

)(1

(L/D)+ γ

)(B.7)

The results for the four constraints described can be seen below for a range of wing loadings,which were plotted on the constraints diagram

W/S (N/m2) 4000 5250 5500 5750 6000 6250 6500 7000 7500Climb to ICA T/W 0.359 0.326 0.322 0.319 0.317 0.316 0.314 0.313 0.314Take-Off T/W 0.18 0.25 0.26 0.27 0.29 0.30 0.31 0.34 0.37Approach (N/m2) W/S 6244 6244 6244 6244 6244 6244 6244 6244 6244Take-off OEI T/W 0.29 0.29 0.29 0.29 0.29 0.29 0.29 0.292 0.292

Table B.1: Constraints analysis detailed results

C Take-off performance

I. Ground distance SG

The ground distance from zero velocity to the take-off velocity was calculated using the followingequation

SG =

∫ VTO

0

V dV

a(C.1)

Where VTO was the take off velocity, given by

VTO = 1.2Vstall = 1.2

√WTO

S

2

ρCLmax(C.2)

Acceleration was found from a free body diagram, given as

a =g

W

[T −D − Ff

]=

g

W

[T −D − µ(WTO − L)

](C.3)

The lift and drag during the ground run are given by

D = 0.5ρV 2S[CD0

+ δCdflap+ δCdgear

+KC2LG

](C.4)

L = 0.5ρV 2SCLG(C.5)

To cross check the landing gear and flap deflection drag coefficients, the following graphs wereused.

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Figure C.1: Change in CDgear with flap deflection (Nicolai, 2011)

Figure C.2: Change in CDflapwith type and deflection of flap Nicolai, 2011

The rolling coeifficient of friction for various take-off and landing distances was found from thefollowing table.

38

Table C.1: Friction coefficient for various surfaces (Nicolai, 2011)

Results obtained during the ground calculations can be seen below.

M Velocity (m/s) Thrust (N) Friction (N) Cd, ground Drag (N) accel. (m/s2) Cum. distance (m) Cum. time (s)0.00 0.00 604257 28694 0.0365 0 2.95 0.00 0.000.01 3.49 597611 28689 0.0365 79 2.92 2.07 1.180.02 6.98 590965 28673 0.0365 317 2.88 8.37 2.380.03 10.47 584319 28647 0.0365 714 2.85 19.00 3.590.04 13.96 577673 28611 0.0365 1269 2.81 34.07 4.810.05 17.45 571027 28564 0.0365 1983 2.77 53.70 6.060.06 20.94 564381 28507 0.0365 2855 2.73 78.03 7.320.07 24.42 557735 28439 0.0365 3886 2.69 107.19 8.590.08 27.91 551089 28361 0.0365 5076 2.65 141.33 9.890.09 31.40 544443 28272 0.0365 6424 2.61 180.61 11.200.10 34.89 537797 28173 0.0365 7931 2.57 225.21 12.540.11 38.38 532173 28064 0.0365 9597 2.54 275.25 13.890.12 41.87 526549 27944 0.0365 11421 2.50 330.87 15.270.13 45.36 520925 27814 0.0365 13404 2.46 392.26 16.670.14 48.85 515301 27673 0.0365 15546 2.42 459.60 18.080.15 52.34 509677 27522 0.0365 17846 2.38 533.13 19.520.16 55.83 504053 27361 0.0365 20304 2.34 613.06 20.990.17 59.32 498429 27189 0.0365 22922 2.30 699.66 22.480.18 62.81 492805 27006 0.0365 25698 2.26 793.18 24.000.19 66.29 487181 26813 0.0365 28632 2.21 893.93 25.540.20 69.78 481557 26610 0.0365 31726 2.17 1002.23 27.120.21 73.27 476704 26397 0.0365 34977 2.13 1118.31 28.730.22 76.76 471852 26173 0.0365 38388 2.09 1242.40 30.370.23 80.25 466999 25938 0.0365 41957 2.05 1374.88 32.040.24 83.09 462147 25740 0.0365 44981 2.01 1489.36 33.43

Table C.2: Ground run during take-off, detailed results

II. Rotation distance SR

The rotation distance lasts for 2 seconds and was given as

SR = 2VTO (C.6)

39

III. Transition distance STR

During transition the aircraft flies at a velocity arc of radius R, where the load factor n was givenas

n = 1 +V 2TO

Rg=

L

W= 1.152 (C.7)

Hence R can be solved by rearraning the top equation

R =VTO

0.152g(C.8)

Assuming the aircraft was an accelerated climb, then the transition distance can be found by,where θCL was the climb gradient.

STR = R sin θCL (C.9)

IV. Climb distance SCL

The climb distance to 35ft was given as

SCL =35 − hTR

tan θCL(C.10)

However, if the transition height was more than 35ft then the climb height was zero.

V. Time during take-off

The time taken during ground was given as

tg =

∫ VTO

0

dV

a(C.11)

The rotation time was assumed to be 2 seconds. The time taken for the transition and climbcan be found by

tTR =STR + (SCL/ cos θCL)

VTO(C.12)

Results during rotation, transition and climb was shown below

40

Rotation phaseVelcoity, rotation (m/s) 85.77Rotation distance (m) 172Time during rotation (s) 2

Transition phaseLoad factor, n 1.152Radius, R (m) 4934Thrust, T (N) 459720Drag, D (N) 198597Climb angle, theta (rad) 0.137Transition distance, Str (m) 673Time during transition (s) 7.85

Climb phaseObstacle height (ft) 35Climb distance (m) -35

Table C.3: Take-off distance, detailed results

D Climb Profile

I. Climb phase

The climb was split into 5000ft step climbs and the rate of climb for each step was found usingthe following equation

R/C = V sin θ = V

[T

W− 1

2ρV 2

(WS

)CD0 −

W

S

2K cos2 θ

ρV 2

](D.1)

However, as the climbs consistedd of small angles it was assumed that the cos θ = 1, which wasa reasonable assumption.

To calculate the time to climb, a numerical solution was used, which was given by

( δh

R/C

)1ststep

=h2 − h1

1/2[(R/C)0 + (R/C)1](D.2)

To calculate the horizontal distance covered and the horizontal velocity, the following equationswere used

Vh =√V 2v − (R/C)2 (D.3)

Sh = Vht+1

2

g

W

[T −D

]t2 (D.4)

41

I. Acceleration

Once a altitude of 1500ft and 10,000ft was reached, the aircraft accelerated to reach the climbspeed. Taking into consideration the loss in thrust and sfc, the distance and time were calculatedusing the following relationship

S =

∫ V2

V1

WV dV

g(T − 12ρV

2SCD)(D.5)

t =

∫ V2

V1

WdV

g(T − 12ρV

2SCD)(D.6)

The take-off, initial climb and acceleration phase of the climb results are shown below

SL 35ft 400ft 1500ft 1500ft accelAircraft weight (kg) 194793 194793 194790 194759 Aircraft weight (kg) 194669Rotation Velocity (m/s) 84 84 84 87 V1, TAS (m/s) 87V,EAS (m/s) 84 84 83 85 V,CAS 129Speed of sound (m/s) 340 340 340 338 Density (kg/m3) 1.172Mach 0.246 0.246 0.246 0.257 V2, TAS (m/s) 132T (N) 228647 228647 228647 221992 Thrust (N) 377935T/W 0.239 0.239 0.239 0.232 Wing area (m2) 306Wing area (m2) 306 306 306 306 Cd0 0.0155W/S (N/m2) 6237 6237 6237 6236 K 0.0413Density (kg/m3) 1.225 1.224 1.2107 1.172 Thrust (N) 377935Cd0 0.0155 0.0155 0.0155 0.0155 time (s) 31K 0.0413 0.0413 0.0413 0.0413 distance (m) 3422R/C (m/s) 14.60 14.59 14.56 14.74 SFC (kg/h/N) 0.0340SFC (kg/hr/N) 0.0320 0.0320 0.0320 0.0318 Fuel (kg) 111Fuel (kg) 0 3 31 90Horizontal velocity (m/s2) 82 82 82 86Distance (m) 0 60 633 2009

Table D.1: Take-off, initial climb and acceleration detailed results

The results from the climb 1,500ft to 10,000ft and acceleration at 10,000ft are shown below

42

1500ft 5000ft 10000ft Acceleration, 10000ftAircraft weight (kg) 194558 194558 194349 Aircraft weight (kg) 194052Rotation Velocity (m/s) 132 139 150 V1, TAS (m/s) 150V,EAS (m/s) 129 129 129 V,CAS (m/s) 155Speed of sound (m/s) 338 334 328 Density (kg/m3) 0.905Mach 0.389 0.415 0.456 V2, TAS (m/s) 195T (N) 311887 145244 128520 Thrust (N) 253560T/W 0.163 0.152 0.135 Wing area (m2) 306Wing area (m2) 306 306 306 Cd0 0.0155W/S (N/m2) 6230 6230 6223 K 0.0413Density (kg/m3) 1.172 1.056 0.905 Thrust (N) 126780Cd0 0.0155 0.0155 0.0155 time (s) 56K 0.0413 0.0413 0.0413 distance (m) 9628R/C (m/s) 15.0 14.2 12.8 SFC (kg/hr/N) 0.0400SFC (kg/hr/N) 0.0362 0.0356 0.0369 Fuel (kg) 158Fuel (kg) 209 297Horizontal velocity (m/s2) 131 138 149Distance (m) 10341 17380

Table D.2: Phase 1 climb detailed results

The results from the climb 10,000ft to 35,000ft are shown below

10000ft 15000ft 20000ft 25000ft 30000ft 35000ftAircraft weight (kg) 193894 193894 193562 193562 193223 192755Rotation Velocity (m/s) 195 200 207 214 223 230V,EAS (m/s) 168 159 151 143 136 128Speed of sound (m/s) 328 322 316 310 303 297Mach 0.594 0.621 0.654 0.691 0.735 0.775T (N) 114121 104763 91007 78392 67814 54626T/W 0.120 0.110 0.096 0.083 0.072 0.058Wing area (m2) 306 306 306 306 306 306W/S (N/m2) 6209 6209 6198 6198 6187 6172Density (kg/m3) 0.905 0.771 0.653 0.549 0.458 0.3796Cd0 0.0155 0.0155 0.0155 0.0155 0.0155 0.0155K 0.0366 0.0413 0.0413 0.0413 0.0413 0.0413R/C (m/s) 11.9 10.8 8.7 6.6 4.7 1.9SFC (kg/hr/N) 0.0432 0.0425 0.0432 0.0446 0.0461 0.0466Fuel (kg) 332 340 339 468 457Horizontal velocity (m/s2) 195 200 207 214 223 230Distance (m) 27315 32727 37853 60846 74880

Table D.3: Phase 2 climb detailed reults

43

Table D.4: Rate of climb at different altitudes

44

E Cruise Profile

I. Acceleration to cruise speed

This method was the same as described in acceleration secction in appendix D, the loss in thrustand sfc due to increasing velocity, was taken into condsideration.

35000ft, accel.Aircraft weight (kg) 192298V1, TAS (m/s) 230V,CAS (m/s) 154.5Density (kg/m3) 0.3796V2, TAS (m/s) 252Thrust (N) 55240Wing area (m2) 306Cd0 0.0155K 0.0413Thrust (N) 110481time (s) 270distance (m) 65124SFC (kg/h/N) 0.0476Fuel (kg) 395

Table E.1: Acceleration to cruise speed results

II. Cruise climb

The master equation was rearranged to give the final weight at the end the range required. AsL/D and V are constant they were left as constants during the integration. This was iterationat 200ft steps consider the effect of changing weight during the cruise and thrust and sfc due toincreasing height.

R =V

C

L

Dlnω (E.1)

The cruise climb results can was shown below,

45

Cruise climbRange (nm) 2097 2097 1048Initial Aircraft weight (kg) 191903 172990 155842Cd0 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413M 0.85 0.85 0.85V,tas (m/s) 252 252 252Cl 0.5087 0.5007 0.4968Cd 0.0261 0.0258 0.0256L/D 19.5 19.4 19.4SFC (kg/h/N) 0.0482 0.0483 0.0484Weight Fraction 1.109 1.110 1.054Final Weight (kg) 172990 155842 147880Fuel (kg) 18913 17148 7962Time (min) 256.7 256.7 128.3density (kg/m3) 0.343 0.314 0.300

Table E.2: Cruise climb detailed results

II. Constant altitude and stepped cruise

The stepped cruise was integration between the cruise at constant altitude and climb mode dis-cussed in appendix ??. During the cruise phase, it would follow the constant altitude relationshipuntil it would climb 2000ft or 1000ft, where it would then follow the climb mode. The constantaltitude relationship was given as

R =

[Vmd

sfc× g

( LD

)max

]2ui

[tan− 1

[ 1

u2i

]− 1

u2iω

]](E.2)

Where(L/D

)max

was found by the following equations

CLopt =

√CDO

K(E.3)( L

D

)max

=CLopt

2CD0(E.4)

The results for the 2,000ft step climb are split into to two tables, constant altitude climbs andcruise climb as shown below

46

Constant altitude cruiseCruise 35000ft Cruise 37000ft Cruise 39000ft

Range (m) 1814960 3825214 3769698Initial Aircraft weight (kg) 191903 181611 163080Cd0 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413V (m/s) 252 248 248density (kg/m3) 0.38 0.348 0.316q (kg/m2) 12079 10705 9720Wing area (m2) 306 306 306SFC (kg/h/N) 0.0487 0.0486 0.0488Final weight (kg) 182671 163617 147040Fuel (kg) 9233 17994 16041Time (min) 120 257 253.3034827

Table E.3: Stepped cruise 2000ft, constant altitude detailed results

Cruise climbClimb 35000ft Climb 37000ft Climb 37000ft Climb 39000ft

Aircraft weight (kg) 182671 182671 163617 163617Velocity (m/s) 249 248 248 248V,EAS (m/s) 484 482 155 155T (N) 55854 50837 50837 46303T/W 0.0623 0.0567 0.0633 0.0577Wing area (m2) 306 306 306 306W/S (N/m2) 5849 5849 5239 5239Density (kg/m) 0.38 0.348 0.348 0.316Cd0 0.0155 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413 0.0413R/C (m/s) 2.7 1.6 2.9 1.8SFC (kg/h/N) 0.0487 0.0486 0.0486 0.0489Fuel (kg) 1060 536Horizontal velocity (m/s) 249 248 248 248Distance (m) 193626 106538Time (sec) 773 427M 0.84 0.84 0.84 0.84

Table E.4: Stepped cruise 2000ft, climb detailed results

The results for the 1,000ft step climb are split into to two tables, constant altitude climbs andcruise climb as shown below

47

Constant altitude cruise35000ft 36000ft 37000ft 38000ft 39000ft 40000ft

Range (m) 963040 1866298 1904000 1909761 1911032 972294Initial Aircraft weight (kg) 191903 186736 177390 168305 159693 151438Cd0 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153K 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413V (m/s) 252 252 252 252 252 252density (kg/m3) 0.38 0.365 0.348 0.332 0.316 0.302q (kg/m3) 12079 11603 11062 10554 10045 9600Wing area (m2) 306 306 306 306 306 306SFC (kg/h/N) 0.0487 0.0483 0.0486 0.0484 0.0488 0.0486Final weight (kg) 186965 177615 168487 159861 151596 147506Fuel (kg) 4939 9121 8904 8443 8097 3933Time (min) 64 123 126 126 126 64

Table E.5: Stepped climb 1000ft, constant altitude cruise detailed results

Cruise climb35000 36000 36000 37000 37000 38000 38000 39000 39000 40000

Aircraft weight (kg) 186965 186965 177615 177615 168487 168487 159861 159861 151596 151596Velocity (m/s) 248 248 248 247 250 251 250 250 250 250V,EAS (m/s) 482 482 155 155 155 155 155 155 155 155T (N) 55854 51079 51079 50837 50837 48570 48570 46303 46303 43877T/W 0.061 0.056 0.059 0.058 0.062 0.059 0.062 0.059 0.062 0.059Wing area (m2) 306 306 306 306 306 306 306 306 306 306W/S (N/m2) 5987 5987 5687 5687 5395 5395 5119 5119 4854 4854Density (kg/m3) 0.380 0.365 0.365 0.348 0.348 0.332 0.332 0.316 0.316 0.302Cd0 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153K 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413R/C (m/s) 2.4 1.3 1.8 1.9 2.5 1.9 2.6 2.0 2.6 2.0SFC (kk/h/N) 0.0487 0.0486 0.0486 0.0486 0.0486 0.0487 0.0487 0.0489 0.0489 0.0489Fuel (kg) 0 229 0 224 0 182 0 168 0 158Horizontal velocity (m/s) 248 248 248 247 250 251 250 250 250 250Distance (m) 0 41262 0 40600 0 34839 0 33568 0 33342Time (s) 0 166 0 164 0 139 0 134 0 133M 0.84 0.84 0.84 0.84 0.85 0.85 0.85 0.85 0.85 0.85

Table E.6: Stepped cruise 1000ft, climb detailed results

48

Constant altitude cruise, 35,000ftRange (m) 3883283 3883283 1941641Initial Aircraft weight (kg) 191903 172667 154696Cd0 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413V (m/s) 252 252 252density (kg/m3) 0.38 0.38 0.38q (kg/m3) 12079 12079 12079Wing area (m2) 306 306 306SFC (kg/h/N) 0.0483 0.0483 0.0483Final weight (kg) 172667 154696 146120Fuel (kg) 19236 17971 8576Time (min) 257 257 128

Table E.7: Constant altitude cruise at 35,000ft detailed results

Constant altitude cruise, 37,000ftRange (m) 3883283 3883283 1941641Initial Aircraft weight (kg) 191903 173130 155703Cd0 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413V (m/s) 252 252 252density (kg/m3) 0.348 0.348 0.348q (kg/m3) 11062 11062 11062Wing area (m2) 306 306 306SFC (kg/h/N) 0.0483 0.0483 0.0483Final weight (kg) 173130 155703 147425Fuel (kg) 18773 17427 8278Time (min) 257 257 128

Table E.8: Constant altitude cruise at 37,000ft detailed results

F Descent Profile

The descent profile was calculated using the same equations as specified in the ??, however theengines were producing residual thrust and negative rate of climb was achieved. The decelerationstage was also calculated using the similar method.

The results for the descent profile, from 40,000ft to 10,000ft are shown below

49

Descent from 40,000ft to 10,000ft40000ft 30000ft 20000ft 10000ft

Cabin pressure altitude at 41000ft (ft) 6000 6000 6000 6000Change in cabin pressure not to exceed (ft/min) 300 300 300 300Minimum time (min) 15 15 15 15Ha (ft) 40000 40000 40000 40000Hb (ft) 1500 1500 1500 1500M 0.84 0.79 0.67 0.55a (m/s) 295 303 316 328Max rate of descent (ft/min) 2567 1925 1925 1925Aircraft weight (kg) 147039 147039 147023 147006Velocity (m/s) 248 240 212 180V,EAS (m/s) 155 155 155 155T (N) 870 3983 4583 4033T/W 0.0012 0.0055 0.0064 0.0056Wing area (m2) 306 306 306 306W/S (N/m2) 4708 4708 4708 4707Density kg/m3) 0.316 0.458 0.653 0.905Cd0 0.0153 0.0153 0.0153 0.0153K 0.0413 0.0413 0.0413 0.0413R/C (m/s) -12.6 -13.0 -12.1 -10.4WFE (kg/h) 265 242 253 319Fuel (kg) 0 16 17 24Horizontal velocity (m/s) 248 240 211 179Distance (m) 0 55598 49764 46620Time (s) 0 238 244 272

Table F.1: Descent to 10,000ft detailed results

50

Descent from 10,000ft to 1,500ft10000ft 5000ft 1500ft

Cabin pressure altitude at 41000ft (ft) 6000 6000 6000Change in cabin pressure not to exceed (ft/min) 300 300 300Minimum time (min) 20 20 20Ha (ft) 40000 40000 40000Hb (ft) 1500 1500 1500M 0.46 0.41 0.39a (m/s) 328 334 338Max rate of descent (ft/min) 1925 1925 1925Aircraft weight (kg) 146978 146978 146978Velocity (m/s) 150 139 132V,EAS (m/s) 129 129 129T (N) -1811 -1296 -1704T/W 0.0025 0.0018 0.0024Wing area (m2) 306 306 306W/S (N/m3) 4706 4706 4706Density (kg/m3) 0.905 1.056 1.172Cd0 0.0153 0.0153 0.0153K 0.0413 0.0413 0.0413R/C (m/s) -7.5 -7.1 -6.6WFE (kg/h) 327 363 396Fuel (kg) 0 21 17Horizontal velocity (m/s) 150 138 131Distance (m) 0 27804 19855Time to descent (s) 0 209 156

Table F.2: Descent to 1,500ft detailed results

G Flight profile

I. Reserves

The reserve flight profile was calculated using the same methods describe for the mission profile,detailed results are shown below.

51

Climb to 20,000ft during reserves1500ft 5000ft 10000ft 20000ft

Aircraft weight (kg) 145074 145074 144932 144733Velocity (m/s) 132 139 150 176V,EAS (m/s) 129 129 129 129Speed of sound (m/s) 338 334 328 316Mach 0.39 0.41 0.46 0.56T (N) 156660 145244 128520 92320T/W 0.220 0.204 0.181 0.130Wing area (m/2 306 306 306 306W/S (N/m2) 4645 4645 4641 4634Density (kg/m3) 1.172 1.056 0.905 0.653Cd0 0.0155 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413 0.0413R/C (m/s) 22.1 21.1 19.2 13.7SFC (kg/h/N) 0.0341 0.0356 0.0369 0.0399Fuel (kg) 0 142 199 189Horizontal velocity (m/s) 133 140 151 177Distance (m) 0 7121 11791 16697

Table G.1: Climb to 20,000ft during reserves

Constant altitude cruise at Vmp

Range 334790Initial Aircraft weight (kg) 144544Cd0 0.0155K 0.0413V (m/s) 190density (kg/m3) 0.653q (kg/m3) 11760Wing area (m2) 306SFC (kg/h/N) 0.0421Final weight (kg) 142894Fuel (kg) 1650Time (min) 29M 0.6

Table G.2: Constant altitude cruise at minimum drag velocity

52

Descent profile to 1,500ft20,000ft 10,000ft 5,000ft 1,500ft

Cabin pressure altitude at 41000ft, (ft) 6000 6000 6000 6000Change in cabin pressure not to exceed (ft/min) 300 300 300 300Minimum time to descent (min) 15 15 15 15Ha 39000 39000 39000 39000Hb 1500 1500 1500 1500M 0.56 0.46 0.41 0.39a (m/s) 316 328 334 338Max rate of descent (ft/min) 2500 2500 2500 2500Aircraft weight (kg) 142894 142894 142860 142839Velocity (m/s) 176 150 139 132V,EAS (m/s) 129 129 129 129T (N) 2144 1811 1296 1704T/W 0.003 0.003 0.002 0.002Wing area (m2) 306 306 306 306W/S (N/m2) 4576 4576 4574 4574Density (kg/m3) 0.653 0.905 1.056 1.172Cd0 0.0155 0.0155 0.0155 0.0155K 0.0413 0.0413 0.0413 0.0413R/C (m/s) -8.7 -7.5 -7.0 -6.6WFE (kg/h) 269 327 363 396Fuel (kg) 0 34 21 17Horizontal velocity (m/s) 177 150 139 132Distance (m) 0 52767 27987 19982

Time to descent (s) 0 375 210 156

Table G.3: Descent profile to 1,500ft during reserves

II. Start up and taxi out

Fuel was approximately equivalent to 9.2 mins of ground idle fuel flow (2.2 mins for start up),adjusted to be consistent with figures shown in performance manuals. A linear correlationbetween ground idle flow and nominal thrust was assumed.

MTOW (kg) 195000

Time (s) 420SLST (lb) 67948a1 75.46b1 0.00225Fuel (kg) 207

Table G.4: Start up and taxi out results

53

III. Approach and landing

The fuel was approximated by the minimum trip fuel descent at mission landing from 20,000ftto 1,500ft.

Approach and landingTime (s) 300 sFuel (kg) 83 kg

Table G.5: Approach and landing results

IV. Taxi in

Taxi in fuel was approximated by 7/9.2 of the start up and taxi out fuel. Taxi in fuel was takenfrom the reserves fuel allowance and was not double accounted in the total fuel, however it wasincluded in the block fuel.

Time (s) 420Fuel (kg) 157

Table G.6: Taxi in results

V. En-route allowance

Approximated by 5$ of the trip fuel.

Time (s) No allowanceFuel (kg) 1803

Table G.7: En-route allowance results

V. Overshoot

Fuel was approximated as 80% of the taxi-off, initial climb to 250kts CAS with the weight atbegining of diversion.

Time (s) No allowanceFuel (kg) 475

Table G.8: Overshoot results

54

H Landing distance

I. Approach distance

Most commercial aircrafts approach at an angle of θ < 3◦, which was cross checked with thefollowing equation

sin θ =1

L/D− T

W(H.1)

Once the flare height and radius were calculated the airborne distance from a 50ft height obstaclewas calculated by

hf = R(1 − cos theta) (H.2)

R =V 2f

0.2g(H.3)

sa =50 − hf

tan θ(H.4)

It was assumed that the flare velocity Vf was equal to 1.23 × the stall velocity Vstall for com-mercial aircraft, which was averaged between the approach speed and the stall speed.

Approach distance ISACl,max 2.7Aircraft weight (kg) 146,878Vstall (m/s) 53Density (kg/m3) 1.225Wing area (m2) 306V,30 (m/s) 69V,TD (m/s) 61Lift (N) 1,440,873Cd0 0.0752K 0.04604Cf 0.015Cg 0.0165D (N) 404078L/D 3.6S,air (m) 291Time (s) 4

Table H.1: Approach distance detailed results

II. Flare distance

The flare angle and approach angle was assumed to be the same and was calculated by thefollowing equation

55

sf = Rsinθ (H.5)

III. Ground distance

During ground run it was assumed the engines were set to flight idle and did not produce anythrust nor thrust reversers were used. Spoilers, flaps and speed brakes were used to increase thedrag to bring the aircraft to a halt as quickly. To calculate the braking distance, integration timesteps were used, using the following equation

SB =W

2g

∫ 0

VTD

d(V 2)

µWL + (ρ/2)SV 2(CD − µCLG)

(H.6)

Braking distance ISAVelocity Cl Cd Drag Friction a Distance Time61 1.366 0.2021 142458 153261 -2.01 60 057 1.366 0.2021 124481 192145 -2.15 168 255 1.366 0.2021 115947 210604 -2.22 217 351 1.366 0.2021 99788 245556 -2.35 306 549 1.366 0.2021 92163 262049 -2.41 346 645 1.366 0.2021 77822 293068 -2.52 419 741 1.366 0.2021 64693 321466 -2.63 482 939 1.366 0.2021 58583 334681 -2.68 511 935 1.366 0.2021 47272 359146 -2.77 562 1131 1.366 0.2021 37173 380990 -2.85 607 1227 1.366 0.2021 28286 400212 -2.92 644 1423 1.366 0.2021 20612 416812 -2.98 676 1519 1.366 0.2021 14149 430790 -3.03 701 1615 1.366 0.2021 8899 442147 -3.07 722 1811 1.366 0.2021 4860 450882 -3.10 736 197 1.366 0.2021 2034 456995 -3.12 746 203 1.366 0.2021 419 460487 -3.14 750 221 1.366 0.2021 67 461250 -3.14 750 22

Table H.2: Ground braking distance detailed results

I Balanced field length

Balance field length with similar equations as described in take-off and landing performanceanalysis. However, the regulated velocities were considered. Results are shown below for theASD and the ACD.

I. ASD

The results for the accelerate continue distance was shown below

56

Thrust V,ef M Thrust,ef Friction, s Friction,ef Cd Drag, s Drag Vref Windmill a start a vref Vref, dist.604257 0 0.00 604257 28694 28694 0.0365 28694 28694 277 3.0 3.0 0604257 5 0.01 594762 28694 28683 0.0365 28694 28854 416 3.0 2.9 4604257 10 0.03 585268 28694 28649 0.0365 28694 29335 554 3.0 2.9 17604257 15 0.04 575774 28694 28593 0.0365 28694 30136 693 3.0 2.8 39604257 20 0.06 566280 28694 28514 0.0365 28694 31257 831 3.0 2.7 71604257 25 0.07 556785 28694 28413 0.0365 28694 32698 970 3.0 2.7 112604257 30 0.09 547291 28694 28289 0.0365 28694 34459 1108 3.0 2.6 164604257 35 0.10 537797 28694 28143 0.0365 28694 36541 1247 3.0 2.6 227604257 40 0.12 526549 28694 27974 0.0365 28694 38943 1386 3.0 2.5 301604257 45 0.13 515301 28694 27782 0.0365 28694 41666 1524 3.0 2.4 387604257 50 0.15 504053 28694 27568 0.0365 28694 44709 1663 3.0 2.4 486604257 55 0.16 492805 28694 27332 0.0365 28694 48072 1801 3.0 2.3 599604257 60 0.18 481557 28694 27073 0.0365 28694 51755 1940 3.0 2.2 728604257 65 0.19 470309 28694 26791 0.0365 28694 55759 2078 3.0 2.1 872604257 70 0.21 481557 28694 26487 0.0365 28694 60083 2217 3.0 2.2 1029

Table I.1: ASD detailed results part 1

a OEI V1 Drag V1 Friction V1 a v1 V first action Cd Drag, action Friction, V1 accel Stop dist Total dist.1.4 1 28984 28693 1.4 4 0.0425 574204 573786 -2.9 3 91.4 6 29370 28676 1.4 9 0.0425 574496 573416 -2.9 14 401.3 11 30073 28636 1.3 14 0.0425 574903 572775 -2.9 33 871.3 16 31093 28574 1.3 19 0.0425 575423 571863 -3.0 61 1511.3 21 32428 28490 1.3 24 0.0425 576055 570680 -3.0 96 2331.3 26 34080 28384 1.3 29 0.0425 576801 569228 -3.0 140 3331.2 31 36047 28255 1.2 34 0.0425 577659 567508 -3.0 192 4511.2 36 38329 28105 1.2 39 0.0425 578628 565519 -3.0 251 5881.1 41 40923 27932 1.1 43 0.0425 579707 563270 -3.0 318 7441.1 46 43832 27737 1.1 48 0.0425 580896 560756 -3.0 393 9201.1 51 47054 27520 1.1 53 0.0425 582195 557978 -3.0 476 11171.0 56 50590 27282 1.0 58 0.0425 583605 554937 -3.0 566 13351.0 61 54438 27021 1.0 63 0.0425 585124 551633 -3.0 664 15760.9 66 58600 26738 0.9 68 0.0425 586752 548068 -3.0 769 18400.9 71 63126 26429 0.9 73 0.0425 588549 544101 -3.0 886 2129

Table I.2: ASD detailed results part 2

II. ACD

The results for the accelerate stop distance was shown below

Thrust,s V,ef M Thrust,ef Friction Vref Cd Dra,s Drag,ef Windmill a start a vref Distance,ef604256.62 0 0.00 604257 28694 0.0365 28694 28694 277 3.0 3.0 0604256.62 5 0.01 594762 28687 0.0365 28694 28859 416 3.0 2.9 4604256.62 10 0.03 585268 28666 0.0365 28694 29352 554 3.0 2.9 17604256.62 15 0.04 575774 28631 0.0365 28694 30174 693 3.0 2.8 39604256.62 20 0.06 566280 28582 0.0365 28694 31324 831 3.0 2.7 71604256.62 25 0.07 556785 28518 0.0365 28694 32803 970 3.0 2.7 112604256.62 30 0.09 547291 28441 0.0365 28694 34611 1108 3.0 2.6 164604256.62 35 0.10 537797 28349 0.0365 28694 36748 1247 3.0 2.6 227604256.62 40 0.12 526549 28244 0.0365 28694 39214 1386 3.0 2.5 301604256.62 45 0.13 515301 28124 0.0365 28694 42008 1524 3.0 2.4 387604256.62 50 0.15 504053 27991 0.0365 28694 45131 1663 3.0 2.4 486604256.62 55 0.16 492805 27843 0.0365 28694 48583 1801 3.0 2.3 600604256.62 60 0.18 481557 27681 0.0365 28694 52363 1940 3.0 2.2 728604256.62 65 0.19 470309 27505 0.0365 28694 56472 2078 3.0 2.1 873604256.62 70 0.21 481557 27315 0.0365 28694 60910 2217 3.0 2.2 1030

Table I.3: ACD detailed results part 1

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a OEI V1 Drag V1 Friction V1 a v1 Cd Friction Vlof Vlof dist. Trans. Dist. Air Dist. Total distance1.4 1 28984 28694 1.4 0.1389 26727 3496 138 224 38591.4 6 29377 28683 1.4 0.1389 26727 3483 138 224 38551.3 11 30095 28658 1.3 0.1389 26727 3450 138 224 38401.3 16 31138 28619 1.3 0.1389 26727 3399 138 224 38161.3 21 32505 28567 1.3 0.1389 26727 3327 138 224 37801.3 26 34196 28500 1.3 0.1389 26727 3234 138 224 37341.2 31 36211 28420 1.2 0.1389 26727 3118 138 224 36751.2 36 38549 28326 1.2 0.1389 26727 2979 138 224 36031.1 41 41208 28218 1.1 0.1389 26727 2813 138 224 35171.1 46 44190 28096 1.1 0.1389 26727 2620 138 224 34141.1 51 47493 27961 1.0 0.1389 26727 2395 138 224 32941.0 56 51118 27811 1.0 0.1389 26727 2137 138 224 31541.0 61 55063 27648 1.0 0.1389 26727 1840 138 224 29910.9 66 59330 27472 0.9 0.1389 26727 1502 138 224 28020.9 71 63971 27279 0.9 0.1389 26727 1123 138 224 2586

Table I.4: ACD detailed results part 2

J Payload - range

MTOW (kg) 195000 MTOW (kg) 195000 OEW (kg) 98200Max Payload (kg) 39170 Max Fuel (kg) 74135 Max Fuel (kg) 74135OEW (kg) 98200 OEW (kg) 98200 TOW (kg) 172335Total fuel available (kg) 57630 Payload available (kg) 22664 Total fuel available (kg) 7413530 min reserve (kg) 5,633 Total fuel available (kg) 74,136 30 min reserve (kg) 5,633

30 min reserve (kg) 5,632Total trip fuel (kg) 51997 Total trip fuel (kg) 6850310% Contingency (kg) 5200 Total trip fuel (kg) 68503 Total trip fuel available (kg) 68503Trip fuel available (kg) 46798 10% Contingency (kg) 6850 Take off and climb (kg) 2835Take off and climb fuel (kg) 2835 Trip fuel available (kg) 61653 Descent fuel (kg) 183Descent and landing fuel (kg) 183 Take off and climb 2835 Trip fuel available 6850Cruise fuel available (kg) 45258 Descent fuel (kg) 183 Cruise fuel available (kg) 58635

Cruise fuel available (kg) 58635Initial cruise weight (kg) 192165 Initial cruise weight (kg) 192165 Initial cruise weight (kg) 169501Final cruise weight (kg) 146907 Final cruise weight (kg) 133530 Final cruise weight (kg) 110866Fuel ratio 1.31 Fuel ratio 1.44 Fuel ratio 1.53

Range (nm) 5145 Range (nm) 6865 Range (nm) 7419Initial Aircraft weight (kg) 191903 Initial Aircraft weight (kg) 192165 Initial Aircraft weight (kg) 169501Cd0 0.0166 Cd0 0.0166 Cd0 0.0166K 0.0366 K 0.0366 K 0.0366V (m/s) 252 V (m/s) 252 V (m/s) 252density (kg/m3) 0.38 density (kg/m3) 0.38 density (kg/m3) 0.38q (kg/m3) 12079 q (kg/m3) 12079 q (kg/m3) 12079Wing area (m2) 306 Wing area (m2) 306 Wing area (m2) 306SFC (kg/h/N) 0.0483 SFC (kg/h/N) 0.0483 SFC (kg/h/N) 0.0483Final weight (kg) 146907 Final weight (kg) 133530 Final weight (kg) 110866Fuel (kg) 45258 Fuel (kg) 58635 Fuel (kg) 58635

Table J.1: Max payload, economy and ferry range detailed results

The payload range diagram was calculated at standard climb and descent procedures at cruisestarting form 35,000ft at 0.85 mach, shown in above table.

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K Performance regulations CS 25

I. One engine inoperative cruise altitude

L Performance regulations

I. CS 25.105 Take-off

(a) The takeoff speeds prescribed by 25.107, the accelerate-stop distance prescribed by 25.109,the takeoff path prescribed by 25.111, the takeoff distance and takeoff run prescribed by 25.113,and the net takeoff flight path prescribed by 25.115, must be determined in the selected con-figuration for takeoff at each weight, altitude, and ambient temperature within the operationallimits selected by the applicant?

(1) In non-icing conditions; and

(b) No takeoff made to determine the data required by this section may require exceptionalpiloting skill or alertness.

(c) The takeoff data must be based on?

(i) Smooth, dry and wet, hard-surfaced runways; and

(d) The takeoff data must include, within the established operational limits of the airplane, thefollowing operational correction factors:

II. CS 25.107 Take-off speeds

(a) V1 must be established in relation to VEF as follows:

(1) VEF was the calibrated airspeed at which the critical engine was assumed to fail. VEF mustbe selected by the applicant, but may not be less than VMCG determined under 25.149(e).

(b) V2MIN, in terms of calibrated airspeed, may not be less than?

(1) 1.13 VSR for?

(i) Two-engine and three-engine turbopropeller and reciprocating engine powered airplanes;and

(c) V2, in terms of calibrated airspeed, must be selected by the applicant to provide at least thegradient of climb required by 25.121(b) but may not be less than?

(d) VMU was the calibrated airspeed at and above which the airplane can safely lift off theground, and con- tinue the takeoff. VMU speeds must be selected by the applicant throughoutthe range of thrust-to-weight ratios to be certificated. These speeds may be established fromfree air data if these data are verified by ground takeoff tests.

(e) VR, in terms of calibrated airspeed, must be selected in accordance with the conditions ofparagraphs (e)(1) through (4) of this section:

(1) VR may not be less than?

(i) V1;

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(ii) 105 percent of VMC;

(f) VLOF was the calibrated airspeed at which the airplane first becomes airborne.

III. CS 25.109 Accelerate stop distance

(a) The accelerate-stop distance on a dry runway was the greater of the following distances:

(1) The sum of the distances necessary to?

(i) Accelerate the airplane from a standing start with all engines operating to VEF for takeofffrom a dry runway;

(ii) Allow the airplane to accelerate from VEF to the highest speed reached during the rejectedtakeoff, assuming the critical engine fails at VEF and the pilot takes the first action to reject thetakeoff at the V1 for takeoff from a dry runway; and

(iii) Come to a full stop on a dry runway from the speed reached as prescribed in paragraph(a)(1)(ii) of this section; plus

(iv) A distance equivalent to 2 seconds at the V1 for takeoff from a dry runway.

(2) The sum of the distances necessary to?

(i) Accelerate the airplane from a standing start with all engines operating to the highest speedreached during the rejected takeoff, assuming the pilot takes the first action to reject the takeoffat the V1 for takeoff from a dry runway; and

(ii) With all engines still operating, come to a full stop on dry runway from the speed reachedas prescribed in paragraph (a)(2)(i) of this section; plus

(iii) A distance equivalent to 2 seconds at the V1 for takeoff from a dry runway.

(b) The accelerate-stop distance on a wet runway was the greater of the following distances:

(1) The accelerate-stop distance on a dry runway determined in accordance with paragraph (a)of this section; or

(2) The accelerate-stop distance determined in accordance with paragraph (a) of this section,except that the runway was wet and the corresponding wet runway values of VEF and V1 areused. In determining the wet runway accelerate-stop distance, the stopping force from the wheelbrakes may never exceed:

(i) The wheel brakes stopping force determined in meeting the requirements of 25.101(i) andparagraph (a) of this section; and

(ii) The force resulting from the wet runway braking coefficient of friction determined in accor-dance with paragraphs (c) or (d) of this section, as applicable, taking into account the distributionof the normal load between braked and unbraked wheels at the most adverse center-of-gravityposition approved for takeoff.

(c) The wet runway braking coefficient of friction for a smooth wet runway was defined as a curveof friction coefficient versus ground speed and must be computed

(d) At the option of the applicant, a higher wet runway braking coefficient of friction may be usedfor runway surfaces that have been grooved or treated with a porous friction course material.

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For grooved and porous friction course runways, the wet runway braking coefficent of frictionwas defined as either:

(1) 70 percent of the dry runway braking coefficient of friction used to determine the dry runwayaccelerate-stop distance; or

(2) The wet runway braking coefficient defined in paragraph (c) of this section, except that aspecific anti-skid system efficiency, if determined, was appropriate for a grooved or porous frictioncourse wet runway, and the maximum tire-to-ground wet runway braking coefficient of frictionwas defined as:

(e) Except as provided in paragraph (f)(1) of this section, means other than wheel brakes maybe used to determine the accelerate-stop distance if that means?

(1) was safe and reliable;

(2) was used so that consistent results can be expected under normal operating conditions;and

(3) was such that exceptional skill was not required to control the airplane.

(g) The landing gear must remain extended throughout the accelerate-stop distance.

(h) If the accelerate-stop distance includes a stopway with surface characteristics substantiallydifferent from those of the runway, the takeoff data must include operational correction fac-tors for the accelerate-stop distance. The correction factors must account for the particularsurface characteristics of the stopway and the variations in these characteristics with seasonalweather conditions (such as temperature, rain, snow, and ice) within the established operationallimits.

(i) A flight test demonstration of the maximum brake kinetic energy accelerate-stop distancemust be conducted with not more than 10 percent of the allowable brake wear range remainingon each of the airplane wheel brakes.

IV. CS 25.111 Take-off path

(a) The takeoff path extends from a standing start to a point in the takeoff at which the airplanewas 1,500 feet above the takeoff surface, or at which the transition from the takeoff to the enroute configuration was completed and VFTO was reached, whichever point was higher. Inaddition?

(1) The takeoff path must be based on the procedures prescribed in 25.101(f);

(2) The airplane must be accelerated on the ground to VEF, at which point the critical enginemust be made inoperative and remain inoperative for the rest of the takeoff; and

(3) After reaching VEF, the airplane must be accelerated to V2.

(b) During the acceleration to speed V2, the nose gear may be raised off the ground at a speednot less than VR. However, landing gear retraction may not be begun until the airplane wasairborne.

(c) During the takeoff path determination in accordance with paragraphs (a) and (b) of thissection?

(1) The slope of the airborne part of the takeoff path must be positive at each point;

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(2) The airplane must reach V2 before it was 35 feet above the takeoff surface and must continueat a speed as close as practical to, but not less than V2, until it was 400 feet above the takeoffsurface;

(3) At each point along the takeoff path, starting at the point at which the airplane reaches 400feet above the takeoff surface, the available gradient of climb may not be less than?

(i) 1.2 percent for two-engine airplanes;

(ii) 1.5 percent for three-engine airplanes; and

(iii) 1.7 percent for four-engine airplanes.

(4) The airplane configuration may not be changed, except for gear retraction and automaticpropeller feathering, and no change in power or thrust that requires action by the pilot may bemade until the airplane was 400 feet above the takeoff surface; and

(5) If 25.105(a)(2) requires the takeoff path to be determined for flight in icing conditions, theairborne part of the takeoff must be based on the airplane drag:

(i) With the most critical of the takeoff ice accretion(s) defined in Appendices C and O of thispart, as applicable, in accordance with 25.21(g), from a height of 35 feet above the takeoff surfaceup to the point where the airplane was 400 feet above the takeoff surface; and

(ii) With the most critical of the final takeoff ice accretion(s) defined in Appendices C and O ofthis part, as applicable, in accordance with 25.21(g), from the point where the airplane was 400feet above the takeoff surface to the end of the takeoff path.

(d) The takeoff path must be determined by a continuous demonstrated takeoff or by synthesisfrom segments. If the takeoff path was determined by the segmental method?

(1) The segments must be clearly defined and must be related to the distinct changes in theconfiguration, power or thrust, and speed;

(2) The weight of the airplane, the configuration, and the power or thrust must be constantthroughout each segment and must correspond to the most critical condition prevailing in thesegment;

(3) The flight path must be based on the airplane’s performance without ground effect; and

(4) The takeoff path data must be checked by continuous demonstrated takeoffs up to the pointat which the airplane was out of ground effect and its speed was stabilized, to ensure that thepath was conservative relative to the continous path.

The airplane was considered to be out of the ground effect when it reaches a height equal to itswing span.

(e) For airplanes equipped with standby power rocket engines, the takeoff path may be determinedin accordance with section II of appendix E.

V. CS 25.113 Take-off distance and run

(a) Takeoff distance on a dry runway was the greater of?

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(1) The horizontal distance along the takeoff path from the start of the takeoff to the pointat which the airplane was 35 feet above the takeoff surface, determined under 25.111 for a dryrunway; or

(2) 115 percent of the horizontal distance along the takeoff path, with all engines operating, fromthe start of the takeoff to the point at which the airplane was 35 feet above the takeoff surface,as determined by a procedure consistent with 25.111.

(b) Takeoff distance on a wet runway was the greater of?

(1) The takeoff distance on a dry runway determined in accordance with paragraph (a) of thissection; or

(2) The horizontal distance along the takeoff path from the start of the takeoff to the point atwhich the airplane was 15 feet above the takeoff surface, achieved in a manner consistent withthe achievement of V2 before reaching 35 feet above the takeoff surface, determined under 25.111for a wet runway.

(c) If the takeoff distance does not include a clearway, the takeoff run was equal to the takeoffdistance. If the takeoff distance includes a clearway?

(1) The takeoff run on a dry runway was the greater of?

(i) The horizontal distance along the takeoff path from the start of the takeoff to a point equidis-tant between the point at which VLOF was reached and the point at which the airplane was 35feet above the takeoff surface, as determined under 25.111 for a dry runway; or

(ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, fromthe start of the takeoff to a point equidistant between the point at which VLOF was reached andthe point at which the airplane was 35 feet above the takeoff surface, determined by a procedureconsistent with 25.111.

(2) The takeoff run on a wet runway was the greater of?

(i) The horizontal distance along the takeoff path from the start of the takeoff to the point atwhich the airplane was 15 feet above the takeoff surface, achieved in a manner consistent withthe achievement of V2 before reaching 35 feet above the takeoff surface, as determined under25.111 for a wet runway; or

(ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, fromthe start of the takeoff to a point equidistant between the point at which VLOF was reached andthe point at which the airplane was 35 feet above the takeoff surface, determined by a procedureconsistent with 25.111.

VI. CS 25.115 Take-off flight path

(a) The takeoff flight path shall be considered to begin 35 feet above the takeoff surface at theend of the takeoff distance determined in accordance with 25.113(a) or (b), as appropriate forthe runway surface condition.

(b) The net takeoff flight path data must be determined so that they represent the actual takeoffflight paths (determined in accordance with 25.111 and with paragraph (a) of this section)reduced at each point by a gradient of climb equal to?

(1) 0.8 percent for two-engine airplanes;

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(c) The prescribed reduction in climb gradient may be applied as an equivalent reduction inacceleration along that part of the takeoff flight path at which the airplane was accelerated inlevel flight.

VIII. CS 25.121 OEI climb

(a) Takeoff; landing gear extended. In the critical takeoff configuration existing along the flightpath (between the points at which the airplane reaches VLOF and at which the landing gear wasfully retracted) and in the configuration used in 25.111 but without ground effect, the steadygradient of climb must be positive for two-engine airplanes, and not less than 0.3 percent forthree-engine airplanes or 0.5 percent for four-engine airplanes, at VLOF and with?

(1) The critical engine inoperative and the remaining engines at the power or thrust availablewhen retraction of the landing gear was begun in accordance with 25.111 unless there was amore critical power operating condition existing later along the flight path but before the pointat which the landing gear was fully retracted; and

(2) The weight equal to the weight existing when retraction of the landing gear was begun,determined under 25.111.

(b) Takeoff; landing gear retracted. In the takeoff configuration existing at the point of the flightpath at which the landing gear was fully retracted, and in the configuration used in 25.111 butwithout ground effect:

(1) The steady gradient of climb may not be less than 2.4 percent for two-engine airplanes, 2.7percent for three-engine airplanes, and 3.0 percent for four-engine airplanes, at V2 with:

(i) The critical engine inoperative, the remaining engines at the takeoff power or thrust availableat the time the landing gear was fully retracted, determined under 25.111, unless there was amore critical power operating condition existing later along the flight path but before the pointwhere the airplane reaches a height of 400 feet above the takeoff surface; and

(ii) The weight equal to the weight existing when the airplane’s landing gear was fully retracted,determined under 25.111.

(2) The requirements of paragraph (b)(1) of this section must be met:

(i) In non-icing conditions; and

(ii) In icing conditions with the most critical of the takeoff ice accretion(s) defined in AppendicesC and O of this part, as applicable, in accordance with 25.21(g), if in the configuration used toshow compliance with 25.121(b) with this takeoff ice accretion:

(A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by morethan the greater of 3 knots CAS or 3 percent of VSR; or

(B) The degradation of the gradient of climb determined in accordance with 25.121(b) was greaterthan one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in25.115(b).

(c) Final takeoff. In the en route configuration at the end of the takeoff path determined inaccordance with 25.111:

(1) The steady gradient of climb may not be less than 1.2 percent for two-engine airplanes, 1.5 per-cent for three-engine airplanes, and 1.7 percent for four-engine airplanes, at VFTO with?

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(i) The critical engine inoperative and the remaining engines at the available maximum contin-uous power or thrust; and

(ii) The weight equal to the weight existing at the end of the takeoff path, determined under25.111.

(2) The requirements of paragraph (c)(1) of this section must be met:

(i) In non-icing conditions; and

(ii) In icing conditions with the most critical of the final takeoff ice accretion(s) defined inAppendices C and O of this part, as applicable, in accordance with 25.21(g), if in the configurationused to show compliance with 25.121(b) with the takeoff ice accretion used to show compliancewith 25.111(c)(5)(i):

(A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by morethan the greater of 3 knots CAS or 3 percent of VSR; or

(B) The degradation of the gradient of climb determined in accordance with 25.121(b) was greaterthan one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in25.115(b).

(d) Approach. In a configuration corresponding to the normal all-engines-operating procedurein which VSR for this configuration does not exceed 110 percent of the VSR for the relatedall-engines-operating landing configuration:

(1) The steady gradient of climb may not be less than 2.1 percent for two-engine airplanes, 2.4percent for three-engine airplanes, and 2.7 percent for four-engine airplanes, with?

(i) The critical engine inoperative, the remaining engines at the go-around power or thrustsetting;

(ii) The maximum landing weight;

(iii) A climb speed established in connection with normal landing procedures, but not exceeding1.4 VSR; and

(iv) Landing gear retracted.

(2) The requirements of paragraph (d)(1) of this section must be met:

(i) In non-icing conditions; and

(ii) In icing conditions with the most critical of the approach ice accretion(s) defined in Appen-dices C and O of this part, as applicable, in accordance with 25.21(g). The climb speed selectedfor non-icing conditions may be used if the climb speed for icing conditions, computed in accor-dance with paragraph (d)(1)(iii) of this section, does not exceed that for non-icing conditions bymore than the greater of 3 knots CAS or 3 percent.

65

M Technical drawing

A leading edge slat was designed shown in figure M.1. This was designed at the two extremeconditions, clean and at take-off and approach configurations.

Figure M.1: Leading edge slat drawing

The leading edge slats are mainly used to improve the aerodynamics and performance charac-teristics of the aircraft. The slats are deployed during landing and take-off to help increase theperformance and retracted during cruise to reduce drag. To deploy the leading edge slat a trackmechanism will be used, which was currently being used on the A320 and the A350. The slatwas guided by rollers in both x and y directions to open the slot and re-energise the airflow. Thismechanism can be seen in figure M.2, used in the A319.

Figure M.2: Leading edge slat track mechanisim

N Aircraft layout

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O Cabin layout

67

P General assembly