Airbus Civil Aircraft Design

CTA1 - GROUP BUSINESS DESIGN PROJECT
PHASE 3 - AIRCRAFT PERFORMANCE
Baba Kakkar
Aerospace Engineering
May 14th 2015

Executive Summary
Since world war II the demand for commercial aircraft has increased considerably, and in
recent years so has the aerospace technology. Therefore, civil aircraft design has become a
very extensive process. The 2015 Airbus design project involved designing a medium-range
airliner 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 up
with an aircraft design called Azure, that meets the Airbus requirements. Azure has a
passenger capacity of 274 in a two class layout for a design range of 5500nm, with a MTOW
of 195 tons and maximum payload of 42.4 tons.
This report has analysed the performance characteristics of the aircraft while ensuring the
operations complied with the aerospace regulations. In order to carry out the required cal-
culations, aerodynamic and propulsive characteristics were needed. To achieve this T0/WT O
and wing loading WT O/S was sized by constructing a constraints diagram against; time to
climb to ICA, cruise at ICA, approach at 145kts or less and take-off at second segment
climb. A T0/WT O and wing loading WT O/S of 0.316 and 6244 (N/m2
) was chosen, which
would improve the aircraft’s fuel efficiency. With a selected wing area and engine size, the
aircraft performance were assessed.
The flight profile provided by Airbus was used to work through the mission phase and
reserves, to ensure the aircraft operations met the Airbus and CS25 requirements. The
climb profile was optimised by climbing at the maximum excess power speed at different
altitudes. 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 the
mach number and the ability of the aircraft to climb to 37,000ft. It was found that the
optimum mach number was 0.86 but this required the aircraft to climb between 37,000ft
and 38,000ft. This was not possible as firstly at end of climb the aircraft would reach its
service ceiling and also it was not able to climb to that altitude. Therefore cruise mach
number of 0.85 and ICA of 35,000ft were chosen. The cruise profile was then optimised by
considering the cruise climb mode and assessing the best weight to change flight to keep
a maximised (L/D). Cruise time, range and fuel burnt achieved were; 10.7 hours, 5215nm
and 44.4 tons at a cruise climb mode respectively. The descent performance was not limited
by engine performance, therefore the fuel economy was not a critical factor as the fuel
consumption was low. The descent fuel, range and time achieved were; 112kg, 135nm and
23 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, which
provided 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 similar
the A330-300 even with a higher MTOW. Airbus requirement of TOFL was also met at
ISA +15◦
. The landing field length was analysed at MLW and ISA SL, which was found
to be 1830m. However, this distance could be reduced by using thrust reversers or drag
parachutes. A balanced field length analysis was assesed to consider the two possibilities
if 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 significant
speeds.
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 as
5400nm, 7117nm and 7670nm respectively. In comparison to the competitor aircrafts, Azure
could fly at a longer economical range. In addition, also concluded from the flight profile
was the relationship between the block fuel at different range and payloads. This was used
as an input to assess how well Azure performed against its competitors at a mission range
of 3000nm. The block fuel and time at 3000nm was found to be 28.1 tons and 6.5 hours. It
was concluded that the Azure performed 5% better than its competitor A330 neo and 11%
better than the B767 re-engined, against DOC.

6
Technical Specification
Operational item’s
Accommodation
Max seat (Single class) 373
Two class seating 274
No. Abreast 1st
; 2nd
; High 6;8;9
Hold Volume (m3
) 163
Mass (Weight) (kg)
Max. take-off (tons) 195
Max. landing (tons) 170
Zero-fuel (tons) 142
Max. Payload (tons) 42.4
Max. fuel payload (tons) 137
Design payload (tons) 42.4
Design fuel load (tons) 53.7
Operational empty (tons) 98.2
Weight Ratios
Max. payload/Max. T/O 0.22
Max. fuel/Max. T/O 0.28
Max. landing/Max. T/O 0.87
Fuel (litres)
Block (tons) 49
Reserve (tons) 5.6
Dimensions
Fuselage
Length (m) 61.3
Height (m) 6.1
Width (m) 6.1
Finess ratio 2.85
Wing
Area (m2
) 306
Span (m) 58
MAC 5.91
Aspect ratio 11
Taper ratio 0.22
Average (t/c) % 0.09
1/4 Chord Sweep 30.7
High Lift Devices
Trailing edge flaps type Single slotted
Flap span/wing span 0.595
Area (m2
) 33.6
Leading edge flaps type slats
Area (m2
) 38.6
Vertical Tail

7
Area (m2
) 47.6
Height (m) 8.86
Aspect ratio 1.65
Taper Ratio 0.35
1/4 Chord sweep 45
Tail arm (m) 29.4
Horizontal Tail
Area (m2
) 80.6
Height (m) 20.7
Aspect ratio 5
Taper Ratio 0.26
1/4 Chord sweep 30
Tail arm (m) 28.7
Undercarriage
Track (m) 10
Wheelbase (m) 25.3
Turning radius (m) 41.27
No. of wheels (nose;main) 2;8
Main wheel diameter (m) 1.05
Main wheel width (m) 0.53
Nacelle
Length (m) 6.38
Max. width (m) 3.58
Spanwise location 9.398
Performance
Landing
Max. wing load (kg/m2
) 6244
Thrust/weight ratio 0.316
Take-oﬀ (m)
ISA sea level 2550
ISA + 15C SL 2681
Landing (m)
ISA sea level 1830
ISA + 15C SL 1894
Speeds (kts/Mach)
Vapp 145
CLmax (T/O) 2.1
CLmax (L/D @ MLM) 2.7
Max cruise
Speed (kts) 489
Altitude (ft) 35,000
Fuel consumption (kg/h) 0.0482
Range (nm)
Max. payload 5396
Design range 5500
Max fuel (+ payload) 7117
Ferry range 7673

Contents
1 Introduction 1
2 Drag polar estimation 2
3 Aircraft sizing 2
3.1 Constraints diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
4 Take-off performance 4
5 Flight profile 7
5.1 Climb performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
5.2 Cruise performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
5.3 Descent performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
5.4 Block performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
5.5 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

9
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 drag
CL Coefficient of lift
CLG
Coefficient of lift in ground at zero angle of attack
D Drag, N
F General force, N
Ff Friction force, N
g Gravitational acceleration, m/s2
htr Transition height, m
K Lift dependant drag factor
L Lift, N
m Mass, kg
M Mach number
n Load factor
p pressure, N/m2
P Power, kW
q Dynamic pressure, N/m2
R Radius, m
S Gross wing area, m2
SG Ground distance m
ST R Transition distance m
SCL Clearance height m
TSL Sea level static thrust
t Time, s, min, hr
u Relative velocity, V/V md
R/C Rate of climb, m/s
V Velocity, kts
VAP P Approach speed, kts
VEF Engine failure recognition speed, kts
VLOF Lift-off speed, kts
Vmd Minimum drag speed, kts
Vmp Minimum power speed, kts
VR Rotational speed, kts
VREF Reference landing speed, kts
V1 Decision speed, kts
V2 Take-off safety speed, kts
W Weight, N
WT O Take-off weight, N

Greek symbols
α Angle of attack, deg, rad; thrust lapse rate
β Weight fraction
γ Climb gradient
µ Runway rolling coefficient of frction
θ Clearance angle, deg, rad
ρ Air density, kg/m3
ω Fuel ratio, W1/W2
Abbreviations
ACD Accelerate continue distance
ASD Accelerate stop distance
BFL Balanced field legnth
CAA Civil airworthiness authority
CAS Calibrated airspeed
CG Centre of gravity
EAS Equivalent airspeed
ICA International cruise altitudet
ISA International standard atmosphere
MLW Maximum landing weight
MTOW Maximum take-off weight
MZFW Maximum zero fuel weight
OEW Operating weight empty
SAR Specific air range
TAS True airspeed
TOFL Take-off field length
TOW Take-off weight
List of Figures
1.1 Blue sky aviation group breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . 1
3.1 Constraints diagram showing design point (solid circle) . . . . . . . . . . . . . . . 3
4.1 Take-off distance breakdown at ISA + 15◦
. . . . . . . . . . . . . . . . . . . . . . 5
4.2 Change in take-off field length with different masses . . . . . . . . . . . . . . . . 7
5.1 Climb performance showing altitude against time, fuel and distance . . . . . . . 8
5.2 Maximum excess power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
5.3 Maximum rate of climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
5.4 TAS and EAS profile during climb . . . . . . . . . . . . . . . . . . . . . . . . . . 11
5.5 Cruise mode profiles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
5.6 L/D against Cl (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
5.7 ML/D optimisation (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . . . 14
5.8 Minimum power cruise climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

12
5.9 Optimisation of flight level during cruise . . . . . . . . . . . . . . . . . . . . . . . 16
5.10 Carpet plot at cruise (Fletcher, 2015) . . . . . . . . . . . . . . . . . . . . . . . . 17
5.11 Range function at constant altitude cruise . . . . . . . . . . . . . . . . . . . . . . 17
5.12 Descent performance showing altitude against time, fuel and distance . . . . . . 18
5.13 TAS and EAS profile during descent . . . . . . . . . . . . . . . . . . . . . . . . . 19
5.14 Block fuel and range at different payloads . . . . . . . . . . . . . . . . . . . . . . 20
5.15 Direct operating cost comparison with competitors (Hassan, 2015) . . . . . . . . 20
5.16 Optimum flight profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
6.1 Breakdown of the landing distance . . . . . . . . . . . . . . . . . . . . . . . . . . 23
7.1 Balanced field length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25
8.1 Payload - Range diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
8.2 Mass breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
A.1 Drag polar during cruise (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . . 33
A.2 Drag polar during take-off (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 33
A.3 Drag polar during landing (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 34
A.4 Drag polar during braking (Dawson, 2015) . . . . . . . . . . . . . . . . . . . . . . 34
C.1 Change in CDgear
with flap deflection (Nicolai, 2011) . . . . . . . . . . . . . . . . 37
C.2 Change in CDflap
with type and deflection of flap Nicolai, 2011 . . . . . . . . . . 37
M.1 Leading edge slat drawing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65
M.2 Leading edge slat track mechanisim . . . . . . . . . . . . . . . . . . . . . . . . . . 65
List of Tables
2.1 Estimation of drag polar during stages of flight (Dawson, 2015) . . . . . . . . . . 2
3.1 Wing loading and thrust to weight ratio comparison . . . . . . . . . . . . . . . . 3
4.1 Comparison of take-off distances . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
4.2 Ground run results breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
4.3 Take-off speeds breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
5.1 Summary of results during climb . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
5.2 Phase 1 climb results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
5.3 Phase 2 climb results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
5.4 Design range breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
5.5 Cruise mode comparison against cruise climb . . . . . . . . . . . . . . . . . . . . 12
5.6 ML/D against mach number . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
5.7 Summary of results during cruise climb . . . . . . . . . . . . . . . . . . . . . . . 14
5.8 Summary of results during stepped cruise . . . . . . . . . . . . . . . . . . . . . . 15
5.9 Summary of results during descent . . . . . . . . . . . . . . . . . . . . . . . . . . 18
5.10 Summary of results for reserve phase . . . . . . . . . . . . . . . . . . . . . . . . . 21
5.11 Fuel and time breakdown of flight profile . . . . . . . . . . . . . . . . . . . . . . . 21
6.1 Comparison of landing distances . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
6.2 Summary of Airborne distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23
6.3 Summary of free roll distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
6.4 Summary of braking distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
7.1 Summary of balanced field length . . . . . . . . . . . . . . . . . . . . . . . . . . 25
8.1 Summary of payload-range diagram . . . . . . . . . . . . . . . . . . . . . . . . . 26
9.1 Airbus design specification comparison . . . . . . . . . . . . . . . . . . . . . . . . 28
10.1 take-off distances at different surface runways . . . . . . . . . . . . . . . . . . . . 28
10.2 OEI climb requirement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

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

1
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 passenger
capacity between 200-300, in a 2 class layout. The main key driver was a 15% reduction of direct
operating cost (DOC) compared to the current 2010 state of the art, B767 re-engined and A330
Neo, with a entry into service at 2025. The Blue sky aviation (CTA1) group was made up of 10
members 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 sky
aviation (CTA1) named ’Azure’ and its economic performance as a transport vehicle, to meet
a design range of 5500nm. To asses the performance of the aircraft, calculation of quantities
were assessed, such as; rate of climb, maximum speed, distance travelled, mass of fuel and
length of runway required for take-off and landing (Mair, 1992). The safety parameters affecting
the performance was also taken into consideration to ensure safe operation handling of the
aircraft.

2
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 drag
count 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 appendix
A.
Take-off Cruise Landing Ground run with spolier
CD0 0.0141 0.0155 0.0752 0.1644
K 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 to
the design, shown in figure 3.1. This was then used to determine the final thrust-to-weight ratio
and the wing loading. The four performance constraints that the aircraft was designed against
were:
1. Approach speed during landing of 145kts or less
2. Take-off from a 3000m field length
3. Climb to ICA with 1.1% climb gradient
4. Cruise at 0.85 mach at ICA
5. Take-off with OEI

3
Figure 3.1: Constraints diagram showing design point (solid circle)
Azure B767-300 B787 A330-300
T0/W0 0.316 0.307 0.273 0.282
W0/S (N/m2
) 6244 6283 7637 5863
S 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 and
increased aircraft efficiency, so the engines were not over sized and the largest possible wing area
was chosen. The design point was chosen to give a small amount of clearway. The results can
be seen in table 3.1, with comparison to other competitor aircrafts. The method and full results
can 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 losses
in thrust were considered and α was found to be 0.8. This was plotted against a range of wing
loadings, which were plotted on the constraints diagram. The aircraft was designed to use single
slotted 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 into
account the drag from the landing gear and flap, but also the windmill drag. OEI climb and
take-off will be discussed in more detail in section 10.

4
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. At
the top of climb the throttle setting was set to RC40 and at cruise the pilot throttled back to
RC35. The thrust lapse rate α at top of climb was 0.181, which dropped to 0.170 at starting
cruise. The fuel fraction was found from the mission profile for both conditions which included
the 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 and
the airborne distance. Conventional aircrafts usually require considerably longer runways which
restrict the number of airports it can access. Therefore, with shorter distances the aircraft will
be more effective economically and operationally. The landing performance will be discussed in
section 6 and the requirements for engine failure will be discussed in section 7.
Azure B767-300 B787 A330-300
Take-off ISA SL (m) 2550 2545 N/A 2320
Take-off ISA+20◦
SL (m) 2725 2850 2678 2680
Take-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 breakdown
of the take-off distances are shown in figure 4.1 at ISA + 15◦
SL. The methods and detailed results
are shown in appendix C.

5
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 of
1518m and took 34 seconds. To calculate the ground distance the model also included the drag
from the landing gear, the flap deflection and the brakes-off friction force. The lift coefficient
at ground run was found from the CL vs α graph, which was 0.731. In addition, the flap and
landing gear drag coefficient were cross checked with the graphs shown in appendix C. The
CDgear predicted by the aerodynamicist was found to be 0.004, however I believe this was very
small compared to a medium/large transporter aircraft. During calculations this change was
found to only decrease the take-off distance by 2%. It was assumed that the flaps were deployed
before the ground run and the CDgear
was constant throughout the ground run. The loss in
thrust and SFC was due to the increased velocity and this was taken into consideration when
integrating between start of ground run to the rotation velocity. The rolling friction coefficient
was found from table D.4, as 0.015.

6
Azure
Weight (kg) 195 tonnes
TE flap deflection 15 deg
Rolling friction coefficient 0.015
Ground 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
0.015
δCDgear
0.004
Air 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 164m
and took 2 seconds. The aircraft rotated on the ground until an angle of attack of 10◦
was reached
which gave a CL of 0.8CLMAX
. According to the GA, a rotation angle of 12◦
was available, hence
there would be no tail strike.
III. Transition distance
The transition distance was the second biggest phase of the take-off distance, which contributed
to 673m and took 8 seconds. During the transition phase the aircraft flew at a constant velocity
arc at radius R. The load factor was found to be 1.152, which gave a radius of 4934m. However
at this radius and takeoff velocity the clearance distance of 35ft was reached and exceeded, hence
the 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 headwind
and tailwind were considered, then it can be seen from the equation in section C, that with a
headwind of 5%, the ground speed would only have needed to be accelerated to VT 0 − VW . This
meant the ground run would have decreased by 10%, and if the tail wind were considered then
the ground run would have increase by 10%. The new total take-off distance would then be
2654m, which would still lie within the Airbus specification. Table 4.3 show a breakdown of the
speeds during the take-off analysis.
Azure
Stall speed VS (kts) 139
Rotation speed VR (kts) 161
Lift 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 with
the maximum take-off weight. Therefore, figure 4.2 shows how the take-off field length varies at
different masses.

7
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 250kts
CAS, which was set by regulations and climb from 10,000ft to 35,000ft at 300kts CAS set by
Airbus specification. However, the second phase of climb was allowed to be changed and was
optimised so better climb rate could be achieved. The method and detailed results can be found
in appendix D.
Azure
Fuel burnt (kg) 2724
Time to climb (min) 22.7
Distance flown (nm) 150
MTOW/ICA (top of climb) 0.986
Table 5.1: Summary of results during climb

8
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 climb
speed. 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 improved
when compared to phase 2B. At the end 10,000ft a (L/D) of 20 was achieved. Table 5.3 highlights
the climb results during this phase.
Azure
Fuel burnt (kg) 664
Rate of climb at 10,000ft (ft/min) 2500
(L/D) at top of climb 20.2
Mach at top of climb 0.46
T0/WT 0 0.135
MTOW/(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 was
designed 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 example
of the results after optimisation at 10,000ft, which were completed for each altitude at 5000ft
step climbs.
Figure 5.2: Maximum excess power
As illustrated on the graph, the best (R/C) were achieved when the there was maximum excess
power between the power available from the thrust and the power required from the drag. The
graph also shows the maximum velocity the aircraft could climb at before more power was
required 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 also
clear from the critical mach number of 0.88. As the thrust varied with airspeed, it was found
that the minimum drag speed did not provide the optimum climb. Figure 5.4 shows how the
TAS and EAS varied during the climb until ICA.

10
Figure 5.3: Maximum rate of climb
Azure
Fuel burnt (kg) 1936
Rate of climb at 35,000ft (ft/min) 375
(L/D) at top of climb 20.2
Mach at top of climb 0.76
T0/WT 0 0.058
MTOW/(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 the
time to climb (Jenkinson, 1999). At higher altitudes the aircraft would not have enough air
density to produce higher thrust, hence the rate of climb would be reduced. This then reflected
on 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 sure
excess fuel was not burnt.
5.2 Cruise performance
To analyse the cruise performance, the cruise range was first identified for the design mission of
5500nm. Using the calculated climb and descent range, the cruise range was found to be 5215nm
as shown in table 5.4. It was decided to cruise at ICA of 35,000ft and the reasons for the choice
will be discussed below.
Azure
Design range (nm) 5500
Climb range (nm) 150
Descent range (nm) 135
Cruise range (nm) 5215
Table 5.4: Design range breakdown

12
The aircraft was capable of cruising in three possible cruise modes, depending on air traffic
control; cruise climb, stepped cruise and cruise at constant altitude. These three modes were
analysed, shown in figure 5.5 and summary of results shown in table 5.5. The methods and
detailed results are shown in appendix E.
Increase from
optimum
Cruise climb fuel (tons) 44.43
Cruise climb time (hr) 10.70
Step 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 mach
number, shown in ﬁgure 5.7. It can be seen that the optimum mach number was between 0.86
and 0.87, however this required a CL between 0.55 and 0.60. This would mean the aircraft would
need to climb between 37,000ft and 38,0000ft, and at the end of cruise the altitude would be
higher then the service ceiling. Therefore the optimum mach number at initial cruise altitude of
35,000ft would be at 0.85. In addition, the aircraft would not be able to produce a positive climb
gradient at 37,000ft, as it was designed on the constraints diagram. This would then require a
higher engine size, which would increase the MTOW. Table 5.6 shows a comparison of cruise
mach 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 A330-300
Cruise speed (kts) 490 489 488 489
ICA (ft) 35000 39000 37000 33000
Cruise 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 the
decrease in aircraft weight as fuel was being burnt. The angle of attack was also kept constant
throughout the cruise, therefore (L/D) was constant at 19.5. The most optimum range during
cruise climb occurred at minimum power velocity, Vmp = 1.136Vmd. However ﬂying at minimum
power velocity was not possible was shown in ﬁgure 5.8. It can be seen that the designed cruise
speed was only 70kts below the optimum. The results calculated during the cruise climb was
integrated at 1000ft steps to allow for the loss in thrust and sfc shown in table 5.7.
Azure
ICA (ft) 35,000
Final cruise altitude (ft) 40,000
L/D 19.5
VT AS (kts) 490
VMP (kts) 560
VMD (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 stepped
cruise mode was iterated to behave in a similar way, as seen in figure 5.5. This was achieved
by setting the range to match equal areas above and below the cruise climb profile. This then
allowed W/ρ to be kept close to the required value. The 2000ft step cruise only used 1.73% more
fuel than the cruise climb, but the time was also increased. This was because when climbing at
2000ft steps, the power available was reduced due to the altitude, and therefore time taken to
climb was longer. During climb the throttle setting was increased to RC40 and due to decrease
in potential energy there was a slight loss in airspeed. The mach number during climb was
also restricted to 0.84, which was assumed to be achieved after throttle change and increase in
altitude. In addition, the 1000ft step climb performed better, as W/ρ was kept closer to required
value. Table 5.8 shows the summary of results.
Azure
ICA (ft) 35,000
Final cruise altitude (ft) 39,000
L/D at 39,000ft 19.8
R/C at 39,000ft (ft/min) 348
Climb 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 to
asses the best weight to change flight (shown in figure 5.9). At a CL of higher than 0.62, the
compressibility 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 and
therefore 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 number
was kept constant. As the cruise speed was greater than the minimum drag speed, then the
decrease in lift coefficient would decrease the drag, meaning thrust will also be decreased. This
meant 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 better
compared 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 compared
to other altitudes. As thrust decreases during the cruise the sfc improves, getting closer to its
low peak. This will not be the case for cruising lower than 37,000ft or higher, as sfc starts to
increase. In addition, figure 5.11 shows the relative velocity (Vmd/V ) against the range for the
constant altitude climb at a fuel fraction of 1.3. Between the initial cruise velocity and the final
cruise cruise velocity, optimum range was achieved. As minimum drag speed reduces during
cruise with aircraft weight the relative velocity increases, reaching the optimum range peak for
best efficiency. Cruising at any higher or lower velocity would decrease the optimum range and
increase 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 from
40,000ft 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 F
were used to calculate the results shown in table 5.9 and figure 5.12, which were integrated at
5,000ft intervals.
Azure
Fuel burnt (kg) 112
Time to descent to 1,500ft (min) 23
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 set
back to RC20, flight idle and the engines produce residual thrust. Therefore, during descent the
fuel economy was not a critical factor as the fuel consumption was low. However, the rate of
change of cabin pressure was to be kept at low value to avoid any problems with passengers on
board. It was agreed that the rate of change of cabin pressure would not exceed 300ft/min at
sea level. With the cabin pressure at 41,000ft set to be equivalent to 6,000ft, the time to descent
to 1,500ft should not have exceeded 15 mins. To avoid not increasing the rate of climb the pilot
may 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 generators
and hydraulic pumps active, so throttle setting could have been chosen to be increased. It was
decided to descend at 300kts CAS as specified by Airbus, but initially at constant mach of 0.84
to avoid the critical mach number. Figure 5.13 shows how the TAS and EAS varied during the
descent 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 on
how much fuel was required for a specific mission. This determined the economic viability of the
aircraft 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, COC
and 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 methods
can be found in appendix G. Table 5.10 shows the summary of the results during the reserve
phase.
Azure
Approach and landing, fuel (kg) 83
Approach and landing, time (min) 5
Taxi in, fuel (kg) 158
Taxi in, fuel (min) 7
En-route allowance, fuel (kg) 1804
Diversion, fuel (kg) 2254
Diversion, time (min) 45
Holding, fuel (kg) 1494
Holding, time (min) 30
Total reserve, fuel (kg) 5634
Total 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 flight
profile was shown in figure 5.16.
Azure
Mission range, fuel (tons) 47.9
Mission range, time (hrs) 11.5
Flight fuel (tons) 48.2
Flight time (hrs) 12
Block fuel (tons) 49
Block 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 the
airborne distance. The ground run distance was the sum of the rolling distance and the braking
distance. The methods described in appendix H have been used to calculate the results shown
in table 6.1. The aircraft mass at the maximum landing weight was used to give the longest
landing distance.
Azure B767-300 B787 A330-300
Landing distance ISA SL (m) 1830 1740 1520 1600
Landing 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 Airbus
requirement.

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 velocity
was 1.3VS. A stall speed of 103kts was achieved from a CL of 2.7, which was optimised by
the aerodynamicist. During calculations, it was assumed that the gradient of the approach was
around 3 degrees, which was similar for commercial aircrafts. The airborne distance contributed
to 21% of the landing distance, which was a total was of 233m and took 6 seconds. To calculate
the air distance the model included the drag from the landing gear and the flap deflection at 10
degrees. The results during the airborne distance can be seen in figure 6.2.
Azure
Weight 147 tonnes
(L/D) 3.6
CDflap
0.015
CDgear
0.004
CLmax 2.7
VS kts 103
V50 135
Air 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 to
flight idle, deployed the spoilers and applied the brakes, which contributed to 10% of landing
distance with a total of 155m. This allowed the rate of descent and airspeed to be reduced to
a safe value. This distance was assumed to last 3 seconds at the touchdown velocity, which was
equal to 1.15VS. Results from the free roll distance are shown in table 6.4.
Azure
Free roll distance 184
CDflap
0.015
CDgear
0.004
CLmax 2.7
VS kts 103
VT D 135
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 of
750m and lasted 22 seconds. The ground lift coefficient CLG
was found from the CL vs α graph,
as 1.366 at the landing configuration.
Azure
CDflap
0.015
CDgear
0.004
CDspoiler
0.006
µ brakes on 0.3
CLG
1.366
(L/D) 6.8
VT D 135
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 and
operationally.

25
7 Balanced field length
The balance field length determined the decision speed V1, from which the aircraft could continue
to accelerate and take-off or stop to a halt when an engine failed. This was calculated by selecting
an engine failure speed VEF from which the pilot had 1 second to make his decision and then 2
seconds to act. As one engine fails at VEF the thrust was reduced and the drag was increased due
to windmilling of one engine. Figure 7.1 and table 7.1 shows the summary of the results.
Figure 7.1: Balanced field length
Azure
VEF (kts) 130
V1 (kts) 132
Vfirstaction (kts) 136
VR (kts) 155
VT O (kts) 162
Balanced 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-range
diagram looked at the balance of range, payload and fuel. Figure 8.1 shows the payload-range
diagram for the aircraft. Main results achieved from this diagram can be seen in table 8.1. This
was achieved with working with the technical integrator and the methods described in appendix
J.
Figure 8.1: Payload - Range diagram
Azure A330-300 B767-300
OWE (tons) 98.2
MTOW (tons) 195
MLW (tons) 170
Max payload (tons) 39.2
Max payload range (nm) 5400 3888 3221
Max economic range (nm) 7117 7056 5354
Ferry 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 found
by the fuel systems manager. An increase in range would require a decrease payload, which would
decrease the drag. However, any range past the economic range will not be economical and DOC
will decrease. If a short mission was required to be completed and the aircraft was needed to
land 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. A
full 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 well
Azure performed against it. All specifications were met.

28
Azure
Passenger 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/MMO
kts CAS /
Mach
360kts /
M=0.89
360kts /
M=0.89
Equivalent Cabin Altitude (at 41000ft) ft 6000 6000
Turn-Around Time mins. Result 52
Airport compatibility limits -
ICAO Code
E
ICAO Code
E
ACN (Flexible B) - 60 60
DOC target $/seat-nm
2010 State
of the art
minus 15%
15%
ETOPS capability (at EIS) mins. 180 293
Expected 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 constantly
referred 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
Azure
Take-off ISA SL (m) concrete 2550
Take-off ISA SL (m) wet concrete 3118
Take-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 ASD
was greater than the distance from standing to highest speed reached during rejected take-off of
1800m.
(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 a
textbook - 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 on
the ground to VEF , at which point the critical engine was made inoperative and after reaching
VEF, the airplane was accelerated to V2
(b) - During the acceleration to speed V2 the landing gear was not retracted at speed of 134kts
which was below VR = 136kts
(c) - The gradient during take-off was positive at each point. At 400ft the gradient with OEI
was 1.8% which was higher than 1.2%
(d) - The take-off path was calculated at segments of time and altitude, where the parameters
were 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-off
distance 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 1,500ft
a 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 climb
gradient was positive.
(b) - At second climb with take-off power with the undercarriage retracted the required climb
gradient was met.
(c) - At final take-off at max continuous power and undercarriage up, the required climb gradient
of 5.8% was met.
Segment 1 Segment 2 Segment 3
Climb 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 appendix
M.
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 were
brought 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 constraints
diagram. This provided adequate information for the group members to perform calculations

31
and finalise the design. Reasonable optimisation during climb and cruise profiles determined
the the most fuel efficient profile the aircraft was capable of. Due to air traffic control is it is
common for aircrafts not to cruise in the most efficient path, hence other cruise modes were also
analysed. It was concluded that other models do incur penalties on fuel burn and time but not
severely.
The performance model also assessed a few WAT analysis, however the model could be improved
by considering an more in depth WAT analysis for each performance characteristic. Airline
companies do not perform missions with maximum take-off or full the mission range, hence
having this analysis could be really useful. The performance model also considered the effects of
the 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 economy
range was also assessed against payload, which helps airline companies manage the payload and
range to maximise profit. On the contrary, after flight test data more accurate and reliable
performance characteristics would be estabilished. Using this performance model and input
from other group members, the Azure was able to perform 11% better then the B767 re-engined
and 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 the
lift 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)

34
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 to
give (MLW/S) (Shevell, 1983):
VAP P ≥ 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 be
included 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:
T0
W0
=
1.44β2
αSGgρCLmax
WT O
S
(B.2)
This equation was altered to give the relevant form to equate the drag polar provided by the
aerodynamicist.
III. Climb to initial cruise altitude
The thrust to weight ratio required to climb to the initial cruise altitude at the design mach
number was calculated by the following master equation:
T0
W0
=
β
α
CD0
(β/q)(WT 0/S)
+ k
β
q
(WT 0/S) +
1
V
dh
dt
(B.3)
aerodynamicist.
β and α are found using the following relationship:
T = αTSL (B.4)
W = βWT O (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 was
calculated by the following master equation:
T0
W0
=
β
α
CD0
(β/q)(WT 0/S)
+ k
β
q
(WT 0/S) (B.6)
aerodynamicist.
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).
T0
W0
=
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 7500
Climb to ICA T/W 0.359 0.326 0.322 0.319 0.317 0.316 0.314 0.313 0.314
Take-Off T/W 0.18 0.25 0.26 0.27 0.29 0.30 0.31 0.34 0.37
Approach (N/m2
) W/S 6244 6244 6244 6244 6244 6244 6244 6244 6244
Take-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 following
equation
SG =
VT O
0
V dV
a
(C.1)
Where VT O was the take off velocity, given by
VT O = 1.2Vstall = 1.2
WT O
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(WT O − L) (C.3)
The lift and drag during the ground run are given by
D = 0.5ρV 2
S CD0
+ δCdflap
+ δCdgear
+ KC2
LG
(C.4)
L = 0.5ρV 2
SCLG
(C.5)
To cross check the landing gear and flap deflection drag coefficients, the following graphs were
used.

37
Figure C.1: Change in CDgear with flap deflection (Nicolai, 2011)
Figure C.2: Change in CDflap
with type and deflection of flap Nicolai, 2011
The rolling coeifficient of friction for various take-off and landing distances was found from the
following 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 0.00
0.01 3.49 597611 28689 0.0365 79 2.92 2.07 1.18
0.02 6.98 590965 28673 0.0365 317 2.88 8.37 2.38
0.03 10.47 584319 28647 0.0365 714 2.85 19.00 3.59
0.04 13.96 577673 28611 0.0365 1269 2.81 34.07 4.81
0.05 17.45 571027 28564 0.0365 1983 2.77 53.70 6.06
0.06 20.94 564381 28507 0.0365 2855 2.73 78.03 7.32
0.07 24.42 557735 28439 0.0365 3886 2.69 107.19 8.59
0.08 27.91 551089 28361 0.0365 5076 2.65 141.33 9.89
0.09 31.40 544443 28272 0.0365 6424 2.61 180.61 11.20
0.10 34.89 537797 28173 0.0365 7931 2.57 225.21 12.54
0.11 38.38 532173 28064 0.0365 9597 2.54 275.25 13.89
0.12 41.87 526549 27944 0.0365 11421 2.50 330.87 15.27
0.13 45.36 520925 27814 0.0365 13404 2.46 392.26 16.67
0.14 48.85 515301 27673 0.0365 15546 2.42 459.60 18.08
0.15 52.34 509677 27522 0.0365 17846 2.38 533.13 19.52
0.16 55.83 504053 27361 0.0365 20304 2.34 613.06 20.99
0.17 59.32 498429 27189 0.0365 22922 2.30 699.66 22.48
0.18 62.81 492805 27006 0.0365 25698 2.26 793.18 24.00
0.19 66.29 487181 26813 0.0365 28632 2.21 893.93 25.54
0.20 69.78 481557 26610 0.0365 31726 2.17 1002.23 27.12
0.21 73.27 476704 26397 0.0365 34977 2.13 1118.31 28.73
0.22 76.76 471852 26173 0.0365 38388 2.09 1242.40 30.37
0.23 80.25 466999 25938 0.0365 41957 2.05 1374.88 32.04
0.24 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 = 2VT O (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 given
as
n = 1 +
V 2
T O
Rg
=
L
W
= 1.152 (C.7)
Hence R can be solved by rearraning the top equation
R =
VT O
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.
ST R = R sin θCL (C.9)
IV. Climb distance SCL
The climb distance to 35ft was given as
SCL =
35 − hT R
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 =
VT O
0
dV
a
(C.11)
The rotation time was assumed to be 2 seconds. The time taken for the transition and climb
can be found by
tT R =
ST R + (SCL/ cos θCL)
VT O
(C.12)
Results during rotation, transition and climb was shown below

40
Rotation phase
Velcoity, rotation (m/s) 85.77
Rotation distance (m) 172
Time during rotation (s) 2
Transition phase
Load factor, n 1.152
Radius, R (m) 4934
Thrust, T (N) 459720
Drag, D (N) 198597
Climb angle, theta (rad) 0.137
Transition distance, Str (m) 673
Time during transition (s) 7.85
Climb phase
Obstacle height (ft) 35
Climb 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 using
the following equation
R/C = V sin θ = V
T
W
−
1
2
ρV 2 W
S
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 was
a 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 equations
were used
Vh = V 2
v − (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 climb
speed. Taking into consideration the loss in thrust and sfc, the distance and time were calculated
using the following relationship
S =
V2
V1
WV dV
g(T − 1
2 ρV 2SCD)
(D.5)
t =
V2
V1
WdV
g(T − 1
2 ρV 2SCD)
(D.6)
The take-oﬀ, initial climb and acceleration phase of the climb results are shown below
SL 35ft 400ft 1500ft 1500ft accel
Aircraft weight (kg) 194793 194793 194790 194759 Aircraft weight (kg) 194669
Rotation Velocity (m/s) 84 84 84 87 V1, TAS (m/s) 87
V,EAS (m/s) 84 84 83 85 V,CAS 129
Speed of sound (m/s) 340 340 340 338 Density (kg/m3
) 1.172
Mach 0.246 0.246 0.246 0.257 V2, TAS (m/s) 132
T (N) 228647 228647 228647 221992 Thrust (N) 377935
T/W 0.239 0.239 0.239 0.232 Wing area (m2
) 306
Wing area (m2
) 306 306 306 306 Cd0 0.0155
W/S (N/m2
) 6237 6237 6237 6236 K 0.0413
Density (kg/m3
) 1.225 1.224 1.2107 1.172 Thrust (N) 377935
Cd0 0.0155 0.0155 0.0155 0.0155 time (s) 31
K 0.0413 0.0413 0.0413 0.0413 distance (m) 3422
R/C (m/s) 14.60 14.59 14.56 14.74 SFC (kg/h/N) 0.0340
SFC (kg/hr/N) 0.0320 0.0320 0.0320 0.0318 Fuel (kg) 111
Fuel (kg) 0 3 31 90
Horizontal velocity (m/s2
) 82 82 82 86
Distance (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, 10000ft
Aircraft weight (kg) 194558 194558 194349 Aircraft weight (kg) 194052
Rotation Velocity (m/s) 132 139 150 V1, TAS (m/s) 150
V,EAS (m/s) 129 129 129 V,CAS (m/s) 155
Speed of sound (m/s) 338 334 328 Density (kg/m3
) 0.905
Mach 0.389 0.415 0.456 V2, TAS (m/s) 195
T (N) 311887 145244 128520 Thrust (N) 253560
T/W 0.163 0.152 0.135 Wing area (m2
) 306
Wing area (m2
) 306 306 306 Cd0 0.0155
W/S (N/m2
) 6230 6230 6223 K 0.0413
Density (kg/m3
) 1.172 1.056 0.905 Thrust (N) 126780
Cd0 0.0155 0.0155 0.0155 time (s) 56
K 0.0413 0.0413 0.0413 distance (m) 9628
R/C (m/s) 15.0 14.2 12.8 SFC (kg/hr/N) 0.0400
SFC (kg/hr/N) 0.0362 0.0356 0.0369 Fuel (kg) 158
Fuel (kg) 209 297
) 131 138 149
Distance (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 35000ft
Aircraft weight (kg) 193894 193894 193562 193562 193223 192755
Rotation Velocity (m/s) 195 200 207 214 223 230
V,EAS (m/s) 168 159 151 143 136 128
Speed of sound (m/s) 328 322 316 310 303 297
Mach 0.594 0.621 0.654 0.691 0.735 0.775
T (N) 114121 104763 91007 78392 67814 54626
T/W 0.120 0.110 0.096 0.083 0.072 0.058
Wing area (m2
) 306 306 306 306 306 306
W/S (N/m2
) 6209 6209 6198 6198 6187 6172
Density (kg/m3
) 0.905 0.771 0.653 0.549 0.458 0.3796
Cd0 0.0155 0.0155 0.0155 0.0155 0.0155 0.0155
K 0.0366 0.0413 0.0413 0.0413 0.0413 0.0413
R/C (m/s) 11.9 10.8 8.7 6.6 4.7 1.9
SFC (kg/hr/N) 0.0432 0.0425 0.0432 0.0446 0.0461 0.0466
Fuel (kg) 332 340 339 468 457
) 195 200 207 214 223 230
Distance (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 thrust
and sfc due to increasing velocity, was taken into condsideration.
35000ft, accel.
Aircraft weight (kg) 192298
V1, TAS (m/s) 230
V,CAS (m/s) 154.5
Density (kg/m3
) 0.3796
V2, TAS (m/s) 252
Thrust (N) 55240
Wing area (m2
) 306
Cd0 0.0155
K 0.0413
Thrust (N) 110481
time (s) 270
distance (m) 65124
SFC (kg/h/N) 0.0476
Fuel (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. As
L/D and V are constant they were left as constants during the integration. This was iteration
at 200ft steps consider the effect of changing weight during the cruise and thrust and sfc due to
increasing height.
R =
V
C
L
D
ln ω (E.1)
The cruise climb results can was shown below,

45
Cruise climb
Range (nm) 2097 2097 1048
Initial Aircraft weight (kg) 191903 172990 155842
Cd0 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413
M 0.85 0.85 0.85
V,tas (m/s) 252 252 252
Cl 0.5087 0.5007 0.4968
Cd 0.0261 0.0258 0.0256
L/D 19.5 19.4 19.4
SFC (kg/h/N) 0.0482 0.0483 0.0484
Weight Fraction 1.109 1.110 1.054
Final Weight (kg) 172990 155842 147880
Fuel (kg) 18913 17148 7962
Time (min) 256.7 256.7 128.3
density (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 relationship
until it would climb 2000ft or 1000ft, where it would then follow the climb mode. The constant
altitude relationship was given as
R =
Vmd
sfc × g
L
D max
2ui tan−
1
1
u2
i
−
1
u2
i ω
(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 and
cruise climb as shown below

46
Constant altitude cruise
Cruise 35000ft Cruise 37000ft Cruise 39000ft
Range (m) 1814960 3825214 3769698
Cd0 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413
V (m/s) 252 248 248
density (kg/m3
) 0.38 0.348 0.316
q (kg/m2
) 12079 10705 9720
Wing area (m2
) 306 306 306
SFC (kg/h/N) 0.0487 0.0486 0.0488
Final weight (kg) 182671 163617 147040
Fuel (kg) 9233 17994 16041
Time (min) 120 257 253.3034827
Table E.3: Stepped cruise 2000ft, constant altitude detailed results
Cruise climb
Climb 35000ft Climb 37000ft Climb 37000ft Climb 39000ft
Aircraft weight (kg) 182671 182671 163617 163617
Velocity (m/s) 249 248 248 248
V,EAS (m/s) 484 482 155 155
T (N) 55854 50837 50837 46303
T/W 0.0623 0.0567 0.0633 0.0577
Wing area (m2
) 306 306 306 306
W/S (N/m2
) 5849 5849 5239 5239
Density (kg/m) 0.38 0.348 0.348 0.316
Cd0 0.0155 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413 0.0413
R/C (m/s) 2.7 1.6 2.9 1.8
SFC (kg/h/N) 0.0487 0.0486 0.0486 0.0489
Fuel (kg) 1060 536
Horizontal velocity (m/s) 249 248 248 248
Distance (m) 193626 106538
Time (sec) 773 427
M 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 and
cruise climb as shown below

47
Constant altitude cruise
35000ft 36000ft 37000ft 38000ft 39000ft 40000ft
Range (m) 963040 1866298 1904000 1909761 1911032 972294
Initial Aircraft weight (kg) 191903 186736 177390 168305 159693 151438
Cd0 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153
K 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413
V (m/s) 252 252 252 252 252 252
density (kg/m3
) 0.38 0.365 0.348 0.332 0.316 0.302
q (kg/m3
) 12079 11603 11062 10554 10045 9600
Wing area (m2
) 306 306 306 306 306 306
SFC (kg/h/N) 0.0487 0.0483 0.0486 0.0484 0.0488 0.0486
Final weight (kg) 186965 177615 168487 159861 151596 147506
Fuel (kg) 4939 9121 8904 8443 8097 3933
Time (min) 64 123 126 126 126 64
Table E.5: Stepped climb 1000ft, constant altitude cruise detailed results
Cruise climb
35000 36000 36000 37000 37000 38000 38000 39000 39000 40000
Aircraft weight (kg) 186965 186965 177615 177615 168487 168487 159861 159861 151596 151596
Velocity (m/s) 248 248 248 247 250 251 250 250 250 250
V,EAS (m/s) 482 482 155 155 155 155 155 155 155 155
T (N) 55854 51079 51079 50837 50837 48570 48570 46303 46303 43877
T/W 0.061 0.056 0.059 0.058 0.062 0.059 0.062 0.059 0.062 0.059
Wing area (m2
) 306 306 306 306 306 306 306 306 306 306
W/S (N/m2
) 5987 5987 5687 5687 5395 5395 5119 5119 4854 4854
Density (kg/m3
) 0.380 0.365 0.365 0.348 0.348 0.332 0.332 0.316 0.316 0.302
Cd0 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153 0.0153
K 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413 0.0413
R/C (m/s) 2.4 1.3 1.8 1.9 2.5 1.9 2.6 2.0 2.6 2.0
SFC (kk/h/N) 0.0487 0.0486 0.0486 0.0486 0.0486 0.0487 0.0487 0.0489 0.0489 0.0489
Fuel (kg) 0 229 0 224 0 182 0 168 0 158
Horizontal velocity (m/s) 248 248 248 247 250 251 250 250 250 250
Distance (m) 0 41262 0 40600 0 34839 0 33568 0 33342
Time (s) 0 166 0 164 0 139 0 134 0 133
M 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, 35,000ft
Range (m) 3883283 3883283 1941641
Cd0 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413
V (m/s) 252 252 252
density (kg/m3
) 0.38 0.38 0.38
q (kg/m3
) 12079 12079 12079
Wing area (m2
) 306 306 306
SFC (kg/h/N) 0.0483 0.0483 0.0483
Final weight (kg) 172667 154696 146120
Fuel (kg) 19236 17971 8576
Time (min) 257 257 128
Table E.7: Constant altitude cruise at 35,000ft detailed results
Constant altitude cruise, 37,000ft
Range (m) 3883283 3883283 1941641
Cd0 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413
V (m/s) 252 252 252
density (kg/m3
) 0.348 0.348 0.348
q (kg/m3
) 11062 11062 11062
Wing area (m2
) 306 306 306
SFC (kg/h/N) 0.0483 0.0483 0.0483
Final weight (kg) 173130 155703 147425
Fuel (kg) 18773 17427 8278
Time (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 the
engines were producing residual thrust and negative rate of climb was achieved. The deceleration
stage 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 10,000ft
40000ft 30000ft 20000ft 10000ft
Cabin pressure altitude at 41000ft (ft) 6000 6000 6000 6000
Change in cabin pressure not to exceed (ft/min) 300 300 300 300
Minimum time (min) 15 15 15 15
Ha (ft) 40000 40000 40000 40000
Hb (ft) 1500 1500 1500 1500
M 0.84 0.79 0.67 0.55
a (m/s) 295 303 316 328
Max rate of descent (ft/min) 2567 1925 1925 1925
Velocity (m/s) 248 240 212 180
V,EAS (m/s) 155 155 155 155
T (N) 870 3983 4583 4033
T/W 0.0012 0.0055 0.0064 0.0056
Wing area (m2
) 306 306 306 306
W/S (N/m2
) 4708 4708 4708 4707
Density kg/m3
) 0.316 0.458 0.653 0.905
Cd0 0.0153 0.0153 0.0153 0.0153
K 0.0413 0.0413 0.0413 0.0413
R/C (m/s) -12.6 -13.0 -12.1 -10.4
WFE (kg/h) 265 242 253 319
Fuel (kg) 0 16 17 24
Distance (m) 0 55598 49764 46620
Time (s) 0 238 244 272
Table F.1: Descent to 10,000ft detailed results

50
Descent from 10,000ft to 1,500ft
10000ft 5000ft 1500ft
Cabin pressure altitude at 41000ft (ft) 6000 6000 6000
Change in cabin pressure not to exceed (ft/min) 300 300 300
Minimum time (min) 20 20 20
Ha (ft) 40000 40000 40000
Hb (ft) 1500 1500 1500
M 0.46 0.41 0.39
a (m/s) 328 334 338
Max rate of descent (ft/min) 1925 1925 1925
Aircraft weight (kg) 146978 146978 146978
Velocity (m/s) 150 139 132
V,EAS (m/s) 129 129 129
T (N) -1811 -1296 -1704
T/W 0.0025 0.0018 0.0024
Wing area (m2
) 306 306 306
W/S (N/m3
) 4706 4706 4706
Density (kg/m3
) 0.905 1.056 1.172
Cd0 0.0153 0.0153 0.0153
K 0.0413 0.0413 0.0413
R/C (m/s) -7.5 -7.1 -6.6
WFE (kg/h) 327 363 396
Fuel (kg) 0 21 17
Horizontal velocity (m/s) 150 138 131
Distance (m) 0 27804 19855
Time 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 reserves
1500ft 5000ft 10000ft 20000ft
Velocity (m/s) 132 139 150 176
V,EAS (m/s) 129 129 129 129
Speed of sound (m/s) 338 334 328 316
Mach 0.39 0.41 0.46 0.56
T (N) 156660 145244 128520 92320
T/W 0.220 0.204 0.181 0.130
Wing area (m/2
306 306 306 306
W/S (N/m2
) 4645 4645 4641 4634
Density (kg/m3
) 1.172 1.056 0.905 0.653
Cd0 0.0155 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413 0.0413
R/C (m/s) 22.1 21.1 19.2 13.7
SFC (kg/h/N) 0.0341 0.0356 0.0369 0.0399
Fuel (kg) 0 142 199 189
Distance (m) 0 7121 11791 16697
Table G.1: Climb to 20,000ft during reserves
Constant altitude cruise at Vmp
Range 334790
Initial Aircraft weight (kg) 144544
Cd0 0.0155
K 0.0413
V (m/s) 190
density (kg/m3
) 0.653
q (kg/m3
) 11760
Wing area (m2
) 306
SFC (kg/h/N) 0.0421
Final weight (kg) 142894
Fuel (kg) 1650
Time (min) 29
M 0.6
Table G.2: Constant altitude cruise at minimum drag velocity

52
Descent profile to 1,500ft
20,000ft 10,000ft 5,000ft 1,500ft
Cabin pressure altitude at 41000ft, (ft) 6000 6000 6000 6000
Change in cabin pressure not to exceed (ft/min) 300 300 300 300
Minimum time to descent (min) 15 15 15 15
Ha 39000 39000 39000 39000
Hb 1500 1500 1500 1500
M 0.56 0.46 0.41 0.39
a (m/s) 316 328 334 338
Max rate of descent (ft/min) 2500 2500 2500 2500
Velocity (m/s) 176 150 139 132
V,EAS (m/s) 129 129 129 129
T (N) 2144 1811 1296 1704
T/W 0.003 0.003 0.002 0.002
Wing area (m2
) 306 306 306 306
W/S (N/m2
) 4576 4576 4574 4574
Density (kg/m3
) 0.653 0.905 1.056 1.172
Cd0 0.0155 0.0155 0.0155 0.0155
K 0.0413 0.0413 0.0413 0.0413
R/C (m/s) -8.7 -7.5 -7.0 -6.6
WFE (kg/h) 269 327 363 396
Fuel (kg) 0 34 21 17
Distance (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 correlation
between ground idle flow and nominal thrust was assumed.
MTOW (kg) 195000
Time (s) 420
SLST (lb) 67948
a1 75.46
b1 0.00225
Fuel (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 20,000ft
to 1,500ft.
Approach and landing
Time (s) 300 s
Fuel (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 taken
from the reserves fuel allowance and was not double accounted in the total fuel, however it was
included in the block fuel.
Time (s) 420
Fuel (kg) 157
Table G.6: Taxi in results
V. En-route allowance
Approximated by 5$ of the trip fuel.
Time (s) No allowance
Fuel (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 at
begining of diversion.
Time (s) No allowance
Fuel (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 the
following 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 obstacle
was calculated by
hf = R(1 − cos theta) (H.2)
R =
V 2
f
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 ISA
Cl,max 2.7
Aircraft weight (kg) 146,878
Vstall (m/s) 53
Density (kg/m3
) 1.225
Wing area (m2
) 306
V,30 (m/s) 69
V,TD (m/s) 61
Lift (N) 1,440,873
Cd0 0.0752
K 0.04604
Cf 0.015
Cg 0.0165
D (N) 404078
L/D 3.6
S,air (m) 291
Time (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 the
following 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 any
thrust nor thrust reversers were used. Spoilers, flaps and speed brakes were used to increase the
drag to bring the aircraft to a halt as quickly. To calculate the braking distance, integration time
steps were used, using the following equation
SB =
W
2g
0
VT D
d(V 2
)
µWL + (ρ/2)SV 2(CD − µCLG
)
(H.6)
Braking distance ISA
Velocity Cl Cd Drag Friction a Distance Time
61 1.366 0.2021 142458 153261 -2.01 60 0
57 1.366 0.2021 124481 192145 -2.15 168 2
55 1.366 0.2021 115947 210604 -2.22 217 3
51 1.366 0.2021 99788 245556 -2.35 306 5
49 1.366 0.2021 92163 262049 -2.41 346 6
45 1.366 0.2021 77822 293068 -2.52 419 7
41 1.366 0.2021 64693 321466 -2.63 482 9
39 1.366 0.2021 58583 334681 -2.68 511 9
35 1.366 0.2021 47272 359146 -2.77 562 11
31 1.366 0.2021 37173 380990 -2.85 607 12
27 1.366 0.2021 28286 400212 -2.92 644 14
23 1.366 0.2021 20612 416812 -2.98 676 15
19 1.366 0.2021 14149 430790 -3.03 701 16
15 1.366 0.2021 8899 442147 -3.07 722 18
11 1.366 0.2021 4860 450882 -3.10 736 19
7 1.366 0.2021 2034 456995 -3.12 746 20
3 1.366 0.2021 419 460487 -3.14 750 22
1 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 performance
analysis. However, the regulated velocities were considered. Results are shown below for the
ASD 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 0
604257 5 0.01 594762 28694 28683 0.0365 28694 28854 416 3.0 2.9 4
604257 10 0.03 585268 28694 28649 0.0365 28694 29335 554 3.0 2.9 17
604257 15 0.04 575774 28694 28593 0.0365 28694 30136 693 3.0 2.8 39
604257 20 0.06 566280 28694 28514 0.0365 28694 31257 831 3.0 2.7 71
604257 25 0.07 556785 28694 28413 0.0365 28694 32698 970 3.0 2.7 112
604257 30 0.09 547291 28694 28289 0.0365 28694 34459 1108 3.0 2.6 164
604257 35 0.10 537797 28694 28143 0.0365 28694 36541 1247 3.0 2.6 227
604257 40 0.12 526549 28694 27974 0.0365 28694 38943 1386 3.0 2.5 301
604257 45 0.13 515301 28694 27782 0.0365 28694 41666 1524 3.0 2.4 387
604257 50 0.15 504053 28694 27568 0.0365 28694 44709 1663 3.0 2.4 486
604257 55 0.16 492805 28694 27332 0.0365 28694 48072 1801 3.0 2.3 599
604257 60 0.18 481557 28694 27073 0.0365 28694 51755 1940 3.0 2.2 728
604257 65 0.19 470309 28694 26791 0.0365 28694 55759 2078 3.0 2.1 872
604257 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 9
1.4 6 29370 28676 1.4 9 0.0425 574496 573416 -2.9 14 40
1.3 11 30073 28636 1.3 14 0.0425 574903 572775 -2.9 33 87
1.3 16 31093 28574 1.3 19 0.0425 575423 571863 -3.0 61 151
1.3 21 32428 28490 1.3 24 0.0425 576055 570680 -3.0 96 233
1.3 26 34080 28384 1.3 29 0.0425 576801 569228 -3.0 140 333
1.2 31 36047 28255 1.2 34 0.0425 577659 567508 -3.0 192 451
1.2 36 38329 28105 1.2 39 0.0425 578628 565519 -3.0 251 588
1.1 41 40923 27932 1.1 43 0.0425 579707 563270 -3.0 318 744
1.1 46 43832 27737 1.1 48 0.0425 580896 560756 -3.0 393 920
1.1 51 47054 27520 1.1 53 0.0425 582195 557978 -3.0 476 1117
1.0 56 50590 27282 1.0 58 0.0425 583605 554937 -3.0 566 1335
1.0 61 54438 27021 1.0 63 0.0425 585124 551633 -3.0 664 1576
0.9 66 58600 26738 0.9 68 0.0425 586752 548068 -3.0 769 1840
0.9 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,ef
604256.62 0 0.00 604257 28694 0.0365 28694 28694 277 3.0 3.0 0
604256.62 5 0.01 594762 28687 0.0365 28694 28859 416 3.0 2.9 4
604256.62 10 0.03 585268 28666 0.0365 28694 29352 554 3.0 2.9 17
604256.62 15 0.04 575774 28631 0.0365 28694 30174 693 3.0 2.8 39
604256.62 20 0.06 566280 28582 0.0365 28694 31324 831 3.0 2.7 71
604256.62 25 0.07 556785 28518 0.0365 28694 32803 970 3.0 2.7 112
604256.62 30 0.09 547291 28441 0.0365 28694 34611 1108 3.0 2.6 164
604256.62 35 0.10 537797 28349 0.0365 28694 36748 1247 3.0 2.6 227
604256.62 40 0.12 526549 28244 0.0365 28694 39214 1386 3.0 2.5 301
604256.62 45 0.13 515301 28124 0.0365 28694 42008 1524 3.0 2.4 387
604256.62 50 0.15 504053 27991 0.0365 28694 45131 1663 3.0 2.4 486
604256.62 55 0.16 492805 27843 0.0365 28694 48583 1801 3.0 2.3 600
604256.62 60 0.18 481557 27681 0.0365 28694 52363 1940 3.0 2.2 728
604256.62 65 0.19 470309 27505 0.0365 28694 56472 2078 3.0 2.1 873
604256.62 70 0.21 481557 27315 0.0365 28694 60910 2217 3.0 2.2 1030
Table I.3: ACD detailed results part 1

57
a OEI V1 Drag V1 Friction V1 a v1 Cd Friction Vlof Vlof dist. Trans. Dist. Air Dist. Total distance
1.4 1 28984 28694 1.4 0.1389 26727 3496 138 224 3859
1.4 6 29377 28683 1.4 0.1389 26727 3483 138 224 3855
1.3 11 30095 28658 1.3 0.1389 26727 3450 138 224 3840
1.3 16 31138 28619 1.3 0.1389 26727 3399 138 224 3816
1.3 21 32505 28567 1.3 0.1389 26727 3327 138 224 3780
1.3 26 34196 28500 1.3 0.1389 26727 3234 138 224 3734
1.2 31 36211 28420 1.2 0.1389 26727 3118 138 224 3675
1.2 36 38549 28326 1.2 0.1389 26727 2979 138 224 3603
1.1 41 41208 28218 1.1 0.1389 26727 2813 138 224 3517
1.1 46 44190 28096 1.1 0.1389 26727 2620 138 224 3414
1.1 51 47493 27961 1.0 0.1389 26727 2395 138 224 3294
1.0 56 51118 27811 1.0 0.1389 26727 2137 138 224 3154
1.0 61 55063 27648 1.0 0.1389 26727 1840 138 224 2991
0.9 66 59330 27472 0.9 0.1389 26727 1502 138 224 2802
0.9 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) 98200
Max Payload (kg) 39170 Max Fuel (kg) 74135 Max Fuel (kg) 74135
OEW (kg) 98200 OEW (kg) 98200 TOW (kg) 172335
Total fuel available (kg) 57630 Payload available (kg) 22664 Total fuel available (kg) 74135
30 min reserve (kg) 5,633 Total fuel available (kg) 74,136 30 min reserve (kg) 5,633
30 min reserve (kg) 5,632
Total trip fuel (kg) 51997 Total trip fuel (kg) 68503
10% Contingency (kg) 5200 Total trip fuel (kg) 68503 Total trip fuel available (kg) 68503
Trip fuel available (kg) 46798 10% Contingency (kg) 6850 Take off and climb (kg) 2835
Take off and climb fuel (kg) 2835 Trip fuel available (kg) 61653 Descent fuel (kg) 183
Descent and landing fuel (kg) 183 Take off and climb 2835 Trip fuel available 6850
Cruise fuel available (kg) 45258 Descent fuel (kg) 183 Cruise fuel available (kg) 58635
Cruise fuel available (kg) 58635
Initial cruise weight (kg) 192165 Initial cruise weight (kg) 192165 Initial cruise weight (kg) 169501
Final cruise weight (kg) 146907 Final cruise weight (kg) 133530 Final cruise weight (kg) 110866
Fuel ratio 1.31 Fuel ratio 1.44 Fuel ratio 1.53
Range (nm) 5145 Range (nm) 6865 Range (nm) 7419
Initial Aircraft weight (kg) 191903 Initial Aircraft weight (kg) 192165 Initial Aircraft weight (kg) 169501
Cd0 0.0166 Cd0 0.0166 Cd0 0.0166
K 0.0366 K 0.0366 K 0.0366
V (m/s) 252 V (m/s) 252 V (m/s) 252
density (kg/m3
) 0.38 density (kg/m3
) 0.38 density (kg/m3
) 0.38
q (kg/m3
) 12079 q (kg/m3
) 12079 q (kg/m3
) 12079
Wing area (m2
) 306 Wing area (m2
) 306 Wing area (m2
) 306
SFC (kg/h/N) 0.0483 SFC (kg/h/N) 0.0483 SFC (kg/h/N) 0.0483
Final weight (kg) 146907 Final weight (kg) 133530 Final weight (kg) 110866
Fuel (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 cruise
starting form 35,000ft at 0.85 mach, shown in above table.

58
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 operational
limits selected by the applicant?
(1) In non-icing conditions; and
(b) No takeoff made to determine the data required by this section may require exceptional
piloting 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, the
following 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 must
be 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 the
gradient 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 the
ground, and continue the takeoff. VMU speeds must be selected by the applicant throughout
the range of thrust-to-weight ratios to be certificated. These speeds may be established from
free 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 of
paragraphs (e)(1) through (4) of this section:
(1) VR may not be less than?
(i) V1;

59
(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 takeoff
from a dry runway;
(ii) Allow the airplane to accelerate from VEF to the highest speed reached during the rejected
takeoff, assuming the critical engine fails at VEF and the pilot takes the first action to reject the
takeoff 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 speed
reached during the rejected takeoff, assuming the pilot takes the first action to reject the takeoff
at 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 reached
as 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 are
used. In determining the wet runway accelerate-stop distance, the stopping force from the wheel
brakes may never exceed:
(i) The wheel brakes stopping force determined in meeting the requirements of 25.101(i) and
paragraph (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 distribution
of the normal load between braked and unbraked wheels at the most adverse center-of-gravity
position approved for takeoff.
(c) The wet runway braking coefficient of friction for a smooth wet runway was defined as a curve
of 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 used
for runway surfaces that have been grooved or treated with a porous friction course material.

60
For grooved and porous friction course runways, the wet runway braking coefficent of friction
was defined as either:
(1) 70 percent of the dry runway braking coefficient of friction used to determine the dry runway
accelerate-stop distance; or
(2) The wet runway braking coefficient defined in paragraph (c) of this section, except that a
specific anti-skid system efficiency, if determined, was appropriate for a grooved or porous friction
course wet runway, and the maximum tire-to-ground wet runway braking coefficient of friction
was defined as:
(e) Except as provided in paragraph (f)(1) of this section, means other than wheel brakes may
be 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 substantially
different 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 particular
surface characteristics of the stopway and the variations in these characteristics with seasonal
weather conditions (such as temperature, rain, snow, and ice) within the established operational
limits.
(i) A flight test demonstration of the maximum brake kinetic energy accelerate-stop distance
must be conducted with not more than 10 percent of the allowable brake wear range remaining
on 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 airplane
was 1,500 feet above the takeoff surface, or at which the transition from the takeoff to the en
route configuration was completed and VFTO was reached, whichever point was higher. In
addition?
(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 engine
must 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 speed
not less than VR. However, landing gear retraction may not be begun until the airplane was
airborne.
(c) During the takeoff path determination in accordance with paragraphs (a) and (b) of this
section?
(1) The slope of the airborne part of the takeoff path must be positive at each point;

61
(2) The airplane must reach V2 before it was 35 feet above the takeoff surface and must continue
at a speed as close as practical to, but not less than V2, until it was 400 feet above the takeoff
surface;
(3) At each point along the takeoff path, starting at the point at which the airplane reaches 400
feet 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 automatic
propeller feathering, and no change in power or thrust that requires action by the pilot may be
made 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, the
airborne 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 this
part, as applicable, in accordance with 25.21(g), from a height of 35 feet above the takeoff surface
up 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 of
this part, as applicable, in accordance with 25.21(g), from the point where the airplane was 400
feet 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 synthesis
from 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 the
configuration, power or thrust, and speed;
(2) The weight of the airplane, the configuration, and the power or thrust must be constant
throughout each segment and must correspond to the most critical condition prevailing in the
segment;
(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 point
at which the airplane was out of ground effect and its speed was stabilized, to ensure that the
path 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 its
wing span.
(e) For airplanes equipped with standby power rocket engines, the takeoff path may be determined
in 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?

62
(1) The horizontal distance along the takeoff path from the start of the takeoff to the point
at which the airplane was 35 feet above the takeoff surface, determined under 25.111 for a dry
runway; or
(2) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from
the 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 this
section; or
(2) The horizontal distance along the takeoff path from the start of the takeoff to the point at
which the airplane was 15 feet above the takeoff surface, achieved in a manner consistent with
the achievement of V2 before reaching 35 feet above the takeoff surface, determined under 25.111
for a wet runway.
(c) If the takeoff distance does not include a clearway, the takeoff run was equal to the takeoff
distance. 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 35
feet 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, from
the start of the takeoff to a point equidistant between the point at which VLOF was reached and
the point at which the airplane was 35 feet above the takeoff surface, determined by a procedure
consistent 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 at
which the airplane was 15 feet above the takeoff surface, achieved in a manner consistent with
the achievement of V2 before reaching 35 feet above the takeoff surface, as determined under
25.111 for a wet runway; or
(ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from
the start of the takeoff to a point equidistant between the point at which VLOF was reached and
the point at which the airplane was 35 feet above the takeoff surface, determined by a procedure
consistent 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 the
end of the takeoff distance determined in accordance with 25.113(a) or (b), as appropriate for
the runway surface condition.
(b) The net takeoff flight path data must be determined so that they represent the actual takeoff
flight 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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