Gulfstream G 200

Aircraft Engineering (AVEN 1920)

Gulfstream G200

By Philip Chu

z3 420180

Table of Contents1. Introduction 1

2. Specification 4

3. Calculation 73.1 Coefficient of Lift (CL) 73.2 Drag 103.3 Coefficient of Drag 113.4 Induced Drag 153.5 Parasitic Drag 173.6 Aspect Ratio 183.7 Wing Loading 193.8 Span Loading 213.9 Power Loading 223.10 Lift-to-drag Ratio 233.11 Thrust-to-weight Ratio 243.12 Maximum G-Loading and Maximum Banking Angle 263.13 Tire Pressure 27

4. Wing 285. Empennage 326. Flight Controls 327. Power Plants 348. Structure 399. Landing Gear 3910. Systems 41

10.1 Environmental Control Systems (ECS) 4110.2 Electrical Systems 4310.3 Rain and Ice Protection Systems 4310.4 Fuel Systems 45

10.5 Avionic Systems 4611. Cabin 4812. References 50

1. Introduction

First introduced in 1999, the Gulfstream G200 is a twin engine business jet

manufactured by Gulfstream Aerospace. It was originally designed by Israel

Aircraft Industries (IAI) and the G200 formerly went by the name of IAI

Galaxy along with Yakolev OKB in a risk-sharing partnership back in the

1980s, with the latter responsible for the design and manufacturing of the

forward fuselage and empennage. Due to delays in production schedule by

the latter, EADS Sogerma was later given the job of manufacturing the

fuselage and empennage.

The purpose of the aircraft was to function as a business jet with

transatlantic crossing capabilities, and was later incorporated into several

military fleets acting as light or VIP transport. 250 G200s were produced

before it ceased in 2011 as it was being succeeded by the improved version

of the G200, the Gulfstream G280.

In the following report, the performance parameters in terms of data

and calculations and the aircraft itself will be analyzed in order to relate

them to the G200’s role as a business jet with transatlantic capabilities.

2. Specifications Gulfstream G200Basic Dimensions

Wing Span

Gross Wing Area

Wing Aspect Ratio

Overall Length

Overall Height

Tail Plane Span

Wheel Track

Wheel Base

17.70 m (Over winglets)

34.3 m2

9.1

18.97 m

6.53 m

6.86 m

3.30 m

7.39 m

Weights

Maximum Ramp Weight

Maximum Take-off Weight

Maximum Landing Weight

Maximum Zero Fuel Weight

Mid-Cruise Weight

Maximum Payload

Payload with Maximum Fuel

Fuel Capacity

16,148 kg

16,080 kg

13,608 kg

10,886 kg

12,247 kg

1,724 kg

181 kg

6,804 kg

Altitudes

Maximum Certified Altitude

Service Ceiling

Cruising Altitude

13,715 m

12,075 m

11,885 m

Speeds

Cruise Speed 494 kt (915 km/h) at FL310

Cruising Mach Number

VMO

MMO

Vs

[mid-cruise weight=12,247 kg]

459 kt (850 km/h) [normal]

0.75 [Long Range]

0.80 [Normal]

310 kt (574 km/h) from S/L to FL100 [IAS]

330 kt (611 km/h) from FL100 to FL 200

[IAS]

360 kt (667 km/h) from FL200 to FL250 [IAS]

0.85

112 kt (208 km/h)

[flaps and gear down, at maximum landing

weight, IAS]

Powerplant

Thrust

By-pass Ratio

Pratt & Whitney Canada PW306A turbofans

26.9 kN each [flat rated]

4.5:1

Tires

Size 26x6.6R14

Range 6,133 km [2 crew + 4 pax at M0.75, NBAA

IFR reserves]

6,469 km [ferry]

G-limit +2.63/-1 [flaps and gear up]

Cabin Dimensions

Length 7.44m [excluding flight deck]

Width

Height

9.30m [including flight deck]

2.18m [at shoulder]

1.73m [at floor]

1.91m

Figure 1. Blueprint of the Gulfstream G200

3. Calculations

*all calculations are done to 3 decimal places

3.1 Co-efficient of lift (CL)

The coefficient of lift (CL) is a co-efficient without dimension relating the

ratios of lift force to the force produced by dynamic pressure multiplied by

area responsible for creating lift. There are, of course, many more CLs as air

density varies along different flight levels, but the ones for take-off, landing

and cruise are the relatively more major ones.

The formula for calculating the co-efficient of lift is given as:

CL=L

12ρ v2S

,

where L is equal to the lift force, ρis the corresponding air density at the

specified altitude, υrefers to the corresponding speed of the aircraft and S is

the area of the aircraft’s wings. In the following calculations, lift is assumed

to be equal to weight, wing area is constant at 34.3 m2 , and the air density

used is 1.2256 kgm-3 at sea level.

3.1.1 Co-efficient of lift during cruise (CL Cruise)

At cruising altitude (11,885 m):

Air density (p) = 1.2256 kgm -3 x 0.2535

Speed (v) = 850 km/h =236.111 ms -1 (3 d.p.)

Weight (W) = 12,247 kg (mid-cruise weight)

Wing area (S) = 34.3 m 2

CLCruise=12,247 kg x9.8

12x (1.2256kgm−3 x 0.2535 ) x ((236.111ms−1)2 ) x (34.3m2)

=0.404 (3 d.p.)

Therefore, the co-efficient of lift during cruise is 0.404.

3.1.2. Co-efficient of lift during take-off (CL TO)

At sea-level:

Air density (p) =1.2256 kgm -3

Speed (v)= 1.15 x Vs = 1.15 x 208 km/h = 66.444 ms -1 (3 d.p.)*

Weight (W)= 16,080 kg (MTOW)

Wing area (S)= 34.3 m 2

CL ¿¿=16080kg x 9.8

12x (1.2256kg m−3 )x ((66.444ms−1)2 ) x (34.3m2)

= 1.698 (3 d.p.)

Therefore, the co-efficient of lift at take-off is 1.698

*Stall speed provided is in terms of MLW, flaps and gear down.

3.1.3. Co-efficient of lift during landing (CL Landing)

At sea-level:

Air density (p) = 1.2256 kgm -3

Speed (v) = 1.3 x Vs = 1.3 x 208 km/h = 75.111 ms -1 (3 d.p.)

Weight(W) = 13,608 kg (MLW)

Wing area (S) =34.3 m 2

CL Landing=13608kg x 9.8

12x (1.2256kgm−3 ) x ((75.111ms−1)2 )x (34.3m2)

= 1.125 (3 d.p.)

Therefore, the co-efficient of lift at landing is 1.125.

3.1.4. Maximum Co-efficient of lift (CL Max)

By varying the angle of attack (α) of the aerofoil, lift can be varied and

hence, CL can also be varied. As the angle of attack increases, CL increases

linearly until maximum possible lift is achieved, and at this point CL achieves

maximum value and is known as CL Max and α reaches the critical angle of

attack. Beyond this point, any further increase in α does not produce any

more additional lift and starts to decrease instead, resulting in what is

known as an aerodynamic stall. The typical critical angle of attack is at

approximately 15 degrees (Figure 2.).

Assuming at sea-level,


Speed (v)= Vs = 208 km/h =57.778 ms -1 (3 d.p.)

Weight (W)= 16,080 kg (MTOW)


CLMax=16080kg x 9.8

12x (1.2256kgm−3 ) x ((57.778ms−1)2 ) x (34.3m2)

=2.246 (3 d.p.)

Therefore, the maximum coefficient of lift is 2.246

3.2 Drag

The total drag of an aircraft moving through air is the sum of its induced

drag and parasitic drag, in which the former is generated as air is re-directed

by the airfoil to generate lift whilst the latter is the drag generated as a

result of an object moving through a fluid. Parasitic drag can be separated

into a few components, namely form drag being the most prominent, skin

Figure 2. Graph illustrating the relationship between CL

and α

friction and interference drag.

As illustrated by Figure 3., induced drag is the greater component of total

drag at lower airspeeds due to the fact that a larger angle of attack is

required to generate lift. As the airspeed increases, induced drag decreases

but form drag increases as air is flowing at relatively faster velocities around

the aircraft, increasing form drag. In order to quantify drag, the co-efficient

of drag can be implemented.

3.3 Co-efficient of Drag (CD)

The co-efficient of drag (CD) is a dimensionless quantity which measures the

drag upon an object moving through a fluid environment, which in this case,

measures the drag of an aircraft moving through air. The formula for CD is:

Figure 3. Graphical representation of the relationship between Total Drag and Airspeed

CD=T

12ρ v2S

where T is equal to thrust, ρis the corresponding air density at the specified

altitude, υrefers to the corresponding speed of the aircraft and S is the area

of the aircraft’s wings. In the following calculations, thrust is assumed to be

equal to drag where thrust is constant, wing area is constant at 34.3 m2 and

the air density used is 1.2256 kgm-3 at sea level. As total drag is the sum of

induced drag and parasitic drag, the co-efficient of drag can also be

expressed as:

CD=CDI+CD0

3.3.1 Co-efficient of drag during cruise (CD Cruise)



Speed (v) = 850 km/h =236.111 ms -1 (3 d.p.)

Thrust = Thrust = 2x 26.9 kN = 53800 N


CDCruise=53800N

12x (1.2256kgm−3 x0.2535 ) x ((236.111ms−1)2 ) x (34.3m2)

= 0.181 (3 d.p.)

Therefore, the co-efficient of drag during cruise is 0.181.

3.3.2. Co-efficient of drag during take-off (CD TO)

At sea-level:

Air density (p) =1.2256 kgm -3

Speed (v)= 1.15 x Vs = 1.15 x 208 km/h = 66.444 ms -1 (3 d.p.)*

Thrust = 2 x 26.9 kN = 53800 N

Wing area (S)= 34.3 m 2

CD¿¿=53800N

12x (1.2256kg m−3 )x ((66.444ms−1)2 ) x (34.3m2)

= 0.580 (3 d.p.)

Therefore, the co-efficient of drag during take-off is 0.580.

*Stall speed provided is in terms of MLW, flaps and gear down.

3.3.3. Co-efficient of drag during landing (CD Landing)

At sea-level:


Speed (v) = 1.3 x Vs = 1.3 x 208 km/h = 75.111 ms -1 (3 d.p.)

Thrust = 2 x 26.9 kN = 53800 N


CD Landing=53800N

12x (1.2256 kgm−3 ) x ((75.111ms−1)2) x (34.3m2)

= 0.454 (3 d.p.)

Therefore, the co-efficient of drag during landing is 0.454.

3.3.4. Maximum co-efficient of drag (CD Max)

Assuming at sea-level,


Speed (v)= Vs = 208 km/h =57.778 ms -1 (3 d.p.)

Thrust = 2 x 26.9 kN = 53800 N


CDMax=53800N

12x (1.2256kgm−3 ) x ((57.778ms−1)2 ) x (34.3m2)

=0.767 (3 d.p.)

Therefore, the maximum coefficient of drag is 0.767.

3.4 Induced Drag

When an airfoil moves through the air, it not only creates lift through

redirecting air, but drag is also created due to a downforce. As induced drag

is related to lift, induced drag increases as the angle of attack increases.

At lower speeds, induced drag tends to be greater due to the larger angle of

attack required to generate lift to compensate for reduced lift generated due

to low airspeed.

3.4.1. Co-efficient of induced drag during cruise (CDI Cruise)

As the co-efficient of induced drag is calculated at cruise conditions, it is the

lesser component of the co-efficient of drag as implied by Figure 3.

The co-efficient of induced drag is a measurement of the drag generated as

a result of lift generation by the airfoil, in which the formula for induced

drag is :

D I=12ρ ν2SC DI

Figure 4. Diagram of airflow and resultant lift and drag of an airfoil

where CDI=CL

2

πeAR and

CL=L

12ρ v2S ,

∴CDI=L2

14ρ2ν2S2πeAR

In the equation, AR is the aspect ratio which will be discussed further on,

and e is the wing span efficiency value which is assumed to be 0.85 in

calculations.

* for accuracy, the value of AR used in the following calculation is not restricted to 3

decimal places.



Speed (v) = 850 km/h =236.111 ms -1 (3 d.p.)

Weight (W) = 12,247 kg (mid-cruise weight)


CL Cruise = 4.04047582 x 10 -3

CD Cruise = 0.181116907

Aspect Ratio (AR) = 9.133819242

During straight and level cruise, lift equals to weight, hence:

CDI=W 2

14ρ2 ν4 S2πeAR

CDI Cruise=(12247kg)2

14x (1.2256kgm−3 x0.2535)2 x (236.111)4 x (34.3m)2 x 0.85 π x (9.133819242)

= 6.969 x 10-5 (3 d.p.)

Therefore, the co-efficient of induced drag during cruise is 6.969 x 10-5.

3.5 Parasitic Drag

Parasitic drag is the larger component of total drag acting on an aircraft at

higher airspeed, and can be separated into a few components, namely form

drag being the most prominent, skin friction and interference drag. Form

drag is the drag generated upon an object moving through a fluid, and as for

skin friction, its occurrence is due to the friction between the flow of air

along the surface of the aircraft. As for interference drag, it mainly occurs

during transonic flow where the range of airspeeds is between Mach 0.8

and 1.2, in which the Gulfstream 200 barely touches the lower limit and

hence is not a major factor contributing towards parasitic drag.

3.5.1. Co-efficient of parasitic drag (CD0)

* for accuracy, the values of CD and CDI used in the following calculation are not restricted

to 3 decimal places.

As the co-efficient of parasitic drag is calculated at cruise conditions, it is the

greater component of the co-efficient of drag as implied by figure b.

Based on the fact that CD=CDI+CD0,

we have CD 0=CD−CDI

CD Cruise = 0.181116907

CDI = 6.969340152 x 10 -5

CD 0=(0.181116901 )−(6.969340152 x10−5 )

= 0.181 (3 d.p.)

Therefore, the co-efficient of parasitic drag is 0.181.

3.6 Aspect Ratio (AR)

The aspect ratio (AR) of an aircraft is one of the indicators of its

performance in terms of maneuverability and efficiency, where the aspect

ratio of an aircraft is given by the formula:

Aspect Ratio=Wingspan2(m2)Wing Area(m2)

.

Generally, aircraft with low aspect ratios then to have greater

maneuverability and are often found on fighter planes due to a higher roll

rate. When compared to a low-AR wing, an equal amount of wing

movement in a high-AR wing due to aileron deflection would have less of a

rolling action on the fuselage due to the relatively longer distance between

the ailerons and the fuselage. This results in a large amount of inertia

needed to be overcome in a maneuver. In terms of efficiency, aircraft with

lower AR compared to one of similar weight but larger AR experiences a

larger induced drag, as a larger downward velocity is needed to lift the

aircraft with the smaller AR. Despite high-AR wings having less induced drag,

they tend to have larger parasitic drag due to a larger wing area and leading-

edge area. Despite so, for aircraft to be as efficient as possible, a higher AR is

preferable as it implies that the wingspan is relatively longer. The effects on

efficiency can be explained through wing-tip vortices, where they only affect

the portion of the wings closet to the wingtips, in which a longer wing would

mean a smaller portion of the wing being affected by the vortices, which

reduces the reduction in lift generated as a result of the vortices, making

flight more fuel efficient in terms of lift.

Also, the section drag co-efficient (Cd) can be inversely correlated to

the chord length to a certain power depending on the airfoil, where a

smaller chord length would result in a larger Cd.

For the Gulfstream 200:

Wingspan = 17.70m

Gross wing area = 34.3m 2

∴Aspect Ratio=(17.70m)2

34.3m2

=9.134 (3 d.p.)

3.7. Wing Loading

Other than aspect ratio, another general indicator of an aircrafts general

maneuverability is its wing loading. The formula for wing loading is given as :

Wing Loading (kgm−2 )= Weight (kg)Wing Area(m2)

Lift is generated by the airfoils of an aircraft due to the motion of air across

the surfaces of the wing. A larger wing area implies that a larger volume of

air is moved, therefore, generates more lift when compared to an aircraft of

similar weight but with a smaller wing area at any velocity. Due to greater

lift generated, aircraft with low wing loading are able to take-off and land at

comparably lower speeds. This can be supported by the formula for the co-

efficient of lift, where:

CL=L

12ρ v2S

, and since L=W based on assumption, CL=

W12ρ v2S

,

in which WS

=12ρ v2CL can be derived where

WS is wing loading.

By further re-arranging the terms,

v2=(2x 9.8)W

SρCL

, in which the relation between speed and wing loading can be

clearly seen. Other than the effect of wing loading on take-off and landing

velocities, it also plays a role in the rate of climb and cruise performance of

an aircraft, in which a lower wing load has a better rate of climb as less

speed is needed to generate the additional lift during climb in relative

terms. During cruise, less thrust is needed to maintain enough lift for level

flight and thus, has a higher cruising efficiency in general. However, the large

wings implied by a small wing load results in greater parasitic drag, and

hence wings with heavier loading are more suited to high speed flight and

maneuvers. In terms of stability, aircraft with higher wing loads then to have

smoother flight as compared to those with lower wing loads.

Gross wing area = 34.3m 2

Weight = 16,080 kg (MTOW)

∴ Wing Loading=16080 kg34.3m2

=468.805 kg m-2 (3 d.p.)

Therefore, the wing loading is 468.805 kg m-2.

3.8. Span Loading

Related to wing loading, span loading is also an indicator of flight stability

and efficiency. The formula for span loading is :

SpanLoading (kgm−1)=Weight (kg)Wingspan(m)

The efficiency behind a smaller span loading can be attributed by the fact

that a longer wingspan results in the wingtip vortices of an aircraft affecting

a relatively smaller portion of the wing as noted in the wing loading section

aforementioned. However, a small span loading renders a plane less

maneuverable due to a larger moment, thus, combat and aerobatic aircraft

generally have a higher span loading, as opposed to long-range airliners and

gliders, where fuel efficiency is more important for the former and greater

lift to sustain flight for a longer time for the latter.

Weight = 16080 kg (MTOW)

Wingspan = 17.70 m

SpanLoading=16080kg17.70m

= 908.475 kg m-1 (3 d.p.)

Therefore, the span loading is 908.475 kg m-1.

3.9. Power Loading (weight-to-power ratio)

The power loading of an aircraft measures actual performance of the

engines and the performance of the aircraft as a whole where the weight is

divided by the power output of the engines. The formula for power loading

is:

Power loading(kgW−1)Weight (kg)Power (W ) ,

where Power= Thrust x Speed.

3.9.1. Power loading during take-off


Thrust = 2x 26.9 kN = 53800 N

Speed = 1.15 x Vs = 1.15x 208km/h =66.444 ms -1 (3 d.p.)

Power Loading= 16080kg

53800N x 66.444ms−1

= 4.498 x 10-3 kg W-1 (3 d.p.)

= 4.498 kg kW-1

Therefore, the power loading during take-off is 4.498 kg kW-1.

3.9.2. Power loading during cruise

Weight = 12,247 kg (mid-cruise weight)

Thrust = 2x 26.9 kN = 53800 N

Speed = 850 km/h =236.111 ms -1 (3 d.p.)


53800x 236.111ms−1

= 9.641 x 10-4 kg W-1 (3 d.p.)

= 0.9641 kg kW-1

Therefore, the power loading during cruise is 0.9641 kg kW-1.

3.9.3. Power loading during landing

Weight = 13,608 kg (MLW)

Thrust = 2x 26.9 kN = 53800 N

Speed = 1.3 x Vs = 1.3 x 208 km/h = 75.111 ms -1 (3 d.p.)


53800x 75.111ms−1

= 3.368 x 10-3 kg W-1 (3 d.p.)

= 3.368 kg kW-1

Therefore, the power loading during landing is 3.368 kg kW-1.

3.10. Lift-to-drag Ratio

The generation of lift produces induced drag, which decreases the efficiency

of which an aircraft is running at. The ratio between lift and drag is an

indicator of the efficiency of an aircraft, in which a higher value is more

favorable as the cost of running that aircraft would be cheaper in terms of

fuel, and aircraft with lower lift-to-drag ratios tend to have better climb

performance.

The lift-to-drag ratio can be determined through a variety of ways such

as flight testing, calculation and wind tunnel tests, in which aircraft

designers try to minimize the lift-to-drag ratios so as to produce an aircraft

with better fuel efficiency. The formula for the lift-to-drag ratio of an aircraft

is :

Lift ¿drag ratio=Lift (N )Drag(N )

As lift is equal to weight during level cruise, and thrust is equal to drag we

have :

Lift ¿drag ratio=9.8 xWeight (kg)

Thrust (N )



Thrust = 2x 26.9 kN = 53800 N

Lift ¿drag ratio=9.8 x12247 kg53800N

=2.231 (3 d.p.)

Therefore, the lift-to-drag ratio is 2.231.

3.11. Thrust-to-weight Ratio

In addition to wing loading, the thrust-to-weight ratio of an aircraft also

serves as a good indicator for aircraft maneuverability. The formula for the

thrust-to-weight ratio of an aircraft is :

Thrust ¿weight ratio=Thrust (N )

9.8 xWeight (kg)

The thrust-to-weight ratio of an aircraft varies along different phases of

flight upon combustion of fuel, resulting in a continual decrease in weight as

flight progresses. Also, thrust varies at different points according to the

throttle setting controlled by the pilot depending on factors such as

airspeed, altitude and temperature. Due to such changes, the thrust-to-

weight ratio quoted is usually the maximum flat thrust divided by the

maximum take-off weight at sea level.

At sea level:


Thrust = 2x 26.9 kN = 53800 N

Thrust ¿weight ratio= 53800N9.8 x16080kg

= 0.341 (3 d.p.)

Therefore, the thrust-to-weight ratio is 0.341.

Due to the fact that lift is equal to weight and thrust is equal to drag in

straight and level flight, the thrust-to-weight ratio during cruise is equal to

the inverse of the lift-to-drag ratio. Hence,

¿



Thrust = 2x 26.9 kN = 53800 N

Thrust ¿weight ratio ¿Cruise=53800N

9.8 x12247kg

= 0.448 (3 d.p.)

Alternatively,

¿

∴ Thrust-to-weight ratioCruise= 0.448 (3 d.p.)

3.12. Maximum G-loading (Maximum load factor/ G-limit) and Maximum

Banking Angle

The G-loading of an aircraft is a measure of the stress subject to the

aircraft’s structure. G-loading is denoted by n, and the formula for n is :

n=Lift (N)

9.8xWeight (kg).

Despite being dimensionless, n is usually referred to in “g”s due to the

correlation between G-loading and the acceleration of gravity felt on board

the aircraft. During straight and level cruise, the G-loading of an aircraft is

1g, in which values of such not equal to one are due to aircraft maneuvers

and/or wind gusts. Positive values for G-loading indicate the aircraft is flying

“the right way up” whilst negative values suggest the opposite, hence the G-

loading for an aircraft and straight and level cruise should be +1g.

The relation between the banking angle and G-loading can be expressed as:

n= 1cosθ

As the maximum G-loading for the Gulfstream 200 is +2.63 with gear and

flaps up, maximum banking angle can be calculated as:

2.63= 1cos θ

θ=cos−1 12.63

= 67.652∘(3 d.p.)

Therefore, the maximum banking angle is 67.652∘.

3.13. Tire pressure

The tire pressure of an aircraft determines which type of runway it can land

on, in which aircraft with higher tire pressures generally require harder

runways capable of withstanding the pressure such that the aircraft can

take-off and land safely without damaging the aircraft or the runway. The

formula for average tire pressure is :

P=9.8 xWeight (lbs)

(number of tires ) x (horizontal cross sectional area of tire )

Based on the tire size of 26x6.6R14 , the horizontal cross-sectional area of

the tire can be calculated, where:

Nominal diameter = 26 inches = 0.6604 m

Nominal section width = 6.6 inches = 0.16764 m


Number of tires on main landing gear = 4

Tire pressure= 9.8 x16080kg4 x (0.6604mx0.16764m)

= 355850.362 Pa

= 51.612 lb/in2

Therefore, the average tire pressure is 51.612 lb/in2.

4. Wings

The Gulfstream G200 has low sweptback wings, which gave the aircraft a

higher center of gravity and increased the wing dihedral, resulting in

increased inherent stability and rolling stability during maneuvers. The

sweptback wings also allowed for the center of gravity to be positioned

towards the aircraft’s aft. The wings of the Gulfstream G200 are based on

its predecessor’s wings, Astra SPX, with modifications such as having a

leading edge sweep of 34° inboard and 25° outboard and integrated

winglets (Figure 5.). The winglets serve to reduce wingtip vortices, hence,

improving fuel efficiency.

Figure 5. Integrated winglet on G200

In addition to the integrated winglets and the inboard leading edge sweep,

Krueger flaps were added as a new feature. Unlike conventional flaps,

Krueger flaps (Figure 6., Figure 10.) are situated on the leading edge of the

airfoil but are not considered as slats due to their different method of

deployment, as they hinge forward from the underside of the airfoil so as to

increase wing camber and lift. Other than the Krueger flaps on the inboard

leading edge, the outboard section is also fitted with slats.

Figure 6. Position and operation of Krueger flap

Figure 7. Location and deployment schematic of Krueger flaps

On the trailing edge of the airfoil are Fowler flaps both situated on the

inboard and outboard sections. Fowler flaps slide backwards before hinging

downwards, which increases the chord length of the airfoil first before

increasing camber. Such flaps do provide some slot effect, in which the

airflow is redirected such that it sticks to the surface of the airfoil, increasing

lift, but the effect is not the main feature of the Fowler flap design.

Figure 8. Schematics of a basic Fowler flap deployment mechanism

Other than the aforementioned slats and flaps,

the upper surface of the airfoil also has four-

segment airbrakes/lift dumpers (Figure 9.).

Figure 10 .Comparison of different flaps

Figure 9. Starboard wing of Gulfstream G200 with winglet and retracted airbrakes visible.

5. Empennage

The Gulfstream G200 has a cruciform tail in which the horizontal stabilizer

intersects the vertical stabilizer above the top of the fuselage (Figure 11.)

Such an arrangement allows the

horizontal stabilizer to be kept

out of the jet engine’s wake

and aids in the avoidance of

interference drag, so as to increase

fuel efficiency. Located on the

empennage are the elevators on

the horizontal stabilizer and the

rudder on the vertical stabilizer.

6. Flight Controls

Flight control surfaces allow a pilot to be able to adjust and control the

altitude of an aircraft, in which the maneuvers can be separated into three

axes of rotation, namely yaw, pitch and roll.

Yaw is controlled by the rudder located on the vertical stabilizer on the

empennage. The rudder is controlled by the rudder pedals in the cockpit

and are operated manually. Regarding directional trim, a small trim tab is

present on the rudder and is operated by two mechanically interconnected

actuators. The rudder bias system is operated by bleed air. As for the

Figure 11. Front view of Gulfstream 200 showing cruciform tail arrangement

maximum amount of deflection for the rudder, it is 20° to the left and right.1

Pitch control is achieved through the hydraulic elevators, located on the

horizontal stabilizer on the empennage, with a maximum movement range

of 27° up and 20° down. Both the rudder and elevator have electronic trim.

Roll control is achieved by the ailerons present on the outboard portion

of the wing, in which the movement range for them is 10° up and 15° down.

The mechanism behind how ailerons aid in roll is based on how the

deployment of ailerons alter the amount of lift on each side of the aircraft

by altering the camber of the airfoil.

The side with the aileron down has increased camber whilst that with the

aileron upwards has decreased camber, in which the side with the

downward aileron generates more lift than that on the other side with the

aileron up. This causes the aircraft to roll towards the side with the aileron

up due to the lift difference generated (Figure 12.). The ailerons themselves

also provide lateral trim.

As for the flaps present in the Gulfstream G200, each wing has inboard

leading-edge Krueger flaps (Figure 7., Figure 13.) with a maximum

deployment of 110° and trailing-edge Fowler flaps located both on the

outboard and the inboard with flap settings at 0, 12, 20 and 40°. The wings

Figure 12. How roll is achieved by ailerons

are also outfitted with outboard leading-edge slats capable of deployment

up to 25°, and each wing is also equipped with four-segment upper surface

airbrakes, which when deployed have a maximum angle of 45°. The

actuators for each control surface are all equipped with torque limiters, in

which an electronic controller stops flap and slat operation when

asymmetrical conditions are created on the wings.

All control surfaces, except for the rudder, are all-hydraulic operated, in

which the hydraulic power required is engine driven at a constant pressure

of 3000 psi. In the event of hydraulic failure, the ailerons and elevators can

be manually operated.

7. Power plant

The Gulfstream G200 is equipped with 2 rear-mounted Pratt & Whitney

Canada PW306A turbofans, each capable of delivering 26.9kN of thrust. The

Figure 13. Schematic of airfoil control surfaces and their respective actuators

power plants are pylon-mounted on each side of the aft fuselage (Figure

14.) Due to the relatively small size of the G200, power plant placement was

impossible under the wings without major structural re-design whilst

maintaining enough wing-engine nacelle and engine nacelle-ground

clearances. Aft fuselage engine placement also brought the advantages of a

greater CL Max as wing-pylon mount and engine exhaust-flap interference

were eliminated, resulting in greater lift at lower speeds. Upon single engine

failure, aft mounted power plants brought less asymmetric yaw to the

aircraft due to a smaller moment of inertia as the engines were close to the

fuselage. In terms of design, aft mounted power plants allowed for the

utilization of shorter landing gear and airstairs as there was adequate

clearance between the ground and the airfoil. There is a quoted factor

regarding why the Gulfstream G200 has aft mounted engines which is

aerodynamically unrelated, and that is the aft mounted power plants

appeared to be aesthetically more appealing.

Despite the advantages regarding efficiency and maneuverability, aft

mounted power plants often incur problems regarding weight distribution.

Figure 14. Rear mounted P ratt & Whitney Canada PW306A turbofan

On an empty G200, the centre of gravity is moved aft to a point where it is

well beyond the point where the centre of gravity for the payload is,

resulting in the need for a larger centre of gravity range. This also leads to

the need for a larger tail so as to compensate for the increased weight. Also,

on wet runways, the wheels may cause water to fly up and be ingested into

the engines, resulting in possible flame-outs, in which special deflectors

have to be installed to prevent such a scenario. Another problem exists

where at very high angles of attack, the nacelle wake blankets the

empennage, often resulting in a deep stall (Figure 15.). Traditionally, aft

mounted engines required a T-tail, but for the G200 a cruciform tail is used

instead. As a preventive measure, a large tail span is usually required,

placing the tail well outboard of the nacelles.

In terms of rolling inertia, asymmetric stall brought on by single engine

failure can result in an excessive roll rate as aft fuselage mounted engines

reduce the rolling moment of inertia when compared to wing mounted

Figure 15. How stall affects planes with T-tails

engines. Last but not least, vibration and noise isolation within the cabin

requires more effort due to the fact that the engines are in close proximity

to the fuselage.

Other than the positioning of the power plants, the mounting of the

power plants to the G200 itself has many factors governing it. The pylons

(Figure 14.) mounting the power plants to the fuselage should be as short as

possible to reduce drag, but long enough in order to avoid any aerodynamic

interference between the engine nacelles, pylons and fuselage of the

aircraft.

Regarding power plant performance itself, the Pratt & Whitney Canada

PW306A turbofans were specifically designed for use on the G200.

The FADEC-equipped turbofan features a 5-stage compressor with a single

centrifugal 4-stage axial with electronically controlled variable Inlet Guide

Vanes and bleed valves. It also has a TALONTM through flow combustor which

allows the aircraft to have reduced nitrous oxide emissions. In terms of

turbines, it consists of a two-stage high pressure turbine and a three-stage

low pressure turbine, in which the combination of the two give the

optimum fuel efficiency. With a by-pass ratio of 4.5, it is considered to be a

Figure 16. Cross-section diagram of a P ratt & Whitney Canada PW306A turbofan

high-bypass turbofan, in which the by-pass ratio is the ratio of air passing

through the engines to air passing around the engines, where higher by-pass

ratios imply a lesser fuel burn and increased fuel efficiency. The Full

Authority Digital Engine Controls (FADEC) for the Pratt & Whitney Canada

PW306A turbofan reduces piloting workload, simplifying operation and

reducing the risk of human error, as FADEC adjusts engine settings in

response to throttle settings and ambient air conditions to provide optimum

output. The total fuel carried in flight for power plant operation is typically

8532 liters, in which 8479 liters is the usable amount.

As for thrust reversers, the system incorporates Nordam nacelles and

hydraulically-actuated thrust reversers (Figure 17.).

The reverse thrust buckets deployed in Figure 17. divert engine exhaust

gases forward, providing a force to decelerate the plane.

Figure 17. Gulfstream G200 with thrust reversers and airbrakes deployed.

8. Structure

The entire fuselage is generally composed of titanium, steel and aluminium

alloy, in which pressure bulkheads are mainly located in the fore and aft of

the baggage compartment and the cabin/cockpit, in which the fuselage fuel

tank is located in between. The aft baggage bulkhead also acts as a support

for the forward engine support beam. For the wings, the main structure of

the airfoil is composed of aluminium alloys and the winglets themselves are

composed of glass-reinforced plastics. The leading edges of the empennage

are also made of composites. As for the auxiliary power unit (APU), it is

housed in the tailcone with a titanium bulkhead.

9. Landing Gear

The landing gear configuration for the Gulfstream G200 is the tricycle

configuration (Figure 19.). Such a configuration has the advantage of being

easier to land as opposed to “tail-dragger”, in which the aircraft has to be

flared before the tailwheel is lowered down onto the runway. Aircraft with

tricycle landing gear configuration are also less vulnerable in a crosswind

landing during the phase where the aircraft is aligned back to the runway

after the crabbing phase just before the nosewheel touches the runway.

The landing gear of the G200 is also retractable, so as to reduce form drag

during cruise as this further streamlines the plane.

The wheels on the main landing gear

are mounted on to a trailing beam

connected to the main strut on the gear

in a pivotal manner, in which the

shock-absorbers are pin-connected,

absorbing beam energy and transmitting

the ground loads to the upper structures

of the gear. Upon gear retraction, it

retracts inwards towards the fuselage

wheel wells(Figure 20.), where the

fuselage door covers the main landing

gear completely so as to reduce drag.

As for the nose landing gear, it

deploys downwards an backwards with

the strut sliding down telescopically from

a rotating tube. The steering angle for

the nose wheel has a maximum value of

Figure 20. Schematic of retractable main landing gear

Fuselage door

Wing door

+60° , in which the steering movement is

transmitted to the wheel axle via torque

links. For towing, an adapter is present on

the strut with an integral safety shear pin.

As mentioned before in the calculations above, the average tire pressure of

the Gulfstream G200 is 51.612 lb/in2.

Runway Surface Pressure withstandable ( lb/in 2 )

Concrete 170-200

Tarmac 70-90

Bad Tarmac 50-70

Hard Grass 45-60

Soft Grass 30-45

Hard-dry Sand 45-60

Wet Sand 25-35

Table 1. Runway surfaces and their corresponding withstandable pressures

From Table 1., the G200 can land on concrete, tarmac and bad tarmac

runways without much problem. This make it versatile in landing and take-

off from a wide array of runways, making it suitable for the role of a private

jet as the versatility allows for flexibility to land in a wide range of airports

around the world.

10. Systems

10.1. Environmental Control System (ECS)

In order to provide a safe and comfortable environment within the cabin for

passengers at high altitudes, the Gulfstream G200 utilizes an environmental

control system. The environmental control system is composed of 5

components, namely: bleed air management, environmental control unit,

temperature control, air distribution and pressurization.

In terms of bleed air management, there are 3 sources of bleed air, which

are the APU, engine low-pressure compressor stage and the engine high-

pressure compressor stage. On the ECS selector in the cockpit, pilots select

what bleed air goes to where, in which normally low-pressure compressor

stage bleed air is used during climb and cruise conditions, whilst high-

pressure compressor stage bleed air is used during high-altitude cruise and

idle descent. Bleed air is extracted from the APU and engines to be used for

air conditioning purposes and pressurization of the cabin. The bleed air

management system draws bleed air through a pre-cooler, pre-cooler by-

pass valve and a thermostat, in which the air is cooled and used for cabin

ventilation. When the low pressure source does not provide minimum cabin

ventilation and cooling, the pre-cooler also provides additional bleed air by

cooling the air from the high-pressure source. In the case of emergency

depressurization, bleed air can be drawn directly from the right engine’s

low-pressure compressor stage into the cabin to provide immediate

pressurization.

The environmental control unit acts as a regulator for cabin pressure,

temperature and ventilation, in which air drawn from the bleed air sources

into heat exchangers and circulated either into the cabin or into the turbine

by-pass valve depending on conditions, in which the former occurs in the

case where conditions inside the cabin are not desirable in a sense that they

are different to the setting indicated in the cockpit, and the latter occurs

when cabin conditions match the settings set by the pilot. Temperature

control limits the temperature between 35°F and 160°F to prevent icing and

protect furnishing materials and occupants from excessive heat respectively.

The air distribution system routes air from the cold air plenum at the ECU

outlet through the fairing in the baggage compartment, in which check

valves are installed in the case of rapid cabin decompression. As for

pressurization, the cabin is pressurized to 0.61 bar or 8.8 lb/sq in.

10.2. Electrical Systems

The aircraft DC Electrical Power System (EPS) is a 28Vdc primary power

system, where power is generated by a pair of 28Vdc, 400A generators each

driven by an engine. As for the main batteries, they serve as back-up power

and are rated at 24Vdc 43AH each, and are used to start engines. A third

battery with a rating of 24Vdc 27AH is used as an emergency power source,

and is connected to the other batteries in parallel. In addition to the pair of

generators, a third 28Vdc 400A generator is driven by the APU, in which this

generator operates in parallel with the other batteries and generators, in

which the APU is only started by the right main battery.

As for power distribution, non-essential heavy load energy consumers are

connected to the main bus whilst those consuming less energy are

connected to the distribution buses. The avionics system is connected to a

separate system, the avionics bus, and is not linked to the main bus or

distribution bus.

10.3. Ice and Rain Protection Systems

The ice and rain protection systems are a group of systems used to protect

the aircraft when operating in rain and ice conditions, namely: airframe de-

icing system, engine de-icing system, ice detection system, probes heat

system, windshield heat system and the windshield wiper system

(Figure 21.).

For pitot probes, static ports, total air temperature probe and the angle

of attack probe, anti-icing is provided by electrical heaters(Figure 21.). For

engine de-icing systems, engine bleed air is distributed inside the leading

edges of the engine nacelle and maybe be further heated with electronic

heaters

(Figure 22.).

Anti-icing of the airfoil and the leading edge of the horizontal stabilizer is

achieved with pneumatically inflated boots.

Regarding the windshield, electrical heating elements installed within the

transparent layers of the windshield provide de-icing, in which constant

windshield temperature is maintained automatically through varying the

electrical power directed to the heating elements, so as to prevent icing of

the windshield and potential cracks from stress brought on to the

windshield through thermal expansion and contraction. As for ice detection,

two detectors are incorporated within the forward fuselage, in which the

presence of icing triggers the detectors, in which the detectors send signals

to the Engine Indicating and Crew Alerting System (EICAS) via the Stall

Protection and Q Feel Computer (SPQC).

In rainy conditions, two-speed windscreen wipers can be activated so as

to remove the rain from the windshields to give clear visibility to the pilots

(Figure 21.)

10.4. Fuel Systems

Fuel is stored in two wing tanks (1,334 litres each), two feed tanks (102 liters

each), a center tank (1,533 liters), a forward tank (1,009 liters) and a

fuselage tank (3,115 liters). Fuel is distributed to each engine via an

independent pressure system, where gravity is the main driving force

distributing fuel between tanks and to the engines. Fuel transfer via

electronic pumps may also occur so as to balance the weights in each

portion of the aircraft so as to maintain a steady center of gravity. As for de-

fueling, there is a single point de-fueling receptacle near the right main

landing gear bay, which is connected to the right main tank.

10.5. Avionics Systems

The avionics system is comprised of an Air Data System, Attitude/Heading

Reference System, Electronic Flight Instrument System, Engine Indication

and Crew Alert System, Flight Control System (Autopilot), Integrated

Avionics Processor System, Maintenance Diagnostics System, Radar Altitude

System, Radio System, Weather Radar System and FMS (Figure 22.).

For its core avionics system, the Gulfstream G200 uses Rockwell Collins Pro

Line 4 Suite Standard.

For communication purposes, the G200 has Dual VHF-422C radios, RTU-

4220 radio tuners, TDR-94D transponders, Baker B1045-F512 audio systems,

triple MagnaStar Flightphones, single Avtech Selcal, Artex ELT and Universal

CVR-30B CVR. Optional add-ons include VHF airborne flight information

Figure 23. Cockpit interior of Gulfstream G200

systems, Bendix/King KHF 950 HF audio systems and 8-channel XM radio.

In terms of flight, the G200 is equipped with the Dual Collins 6100 FMS

with embedded GPS, dual Collins FCC-4005 autopilots, AHS-3000 AHRS,

ADC-850C air data systems, VIR-432 VOR/ILS/GS/markers and DME-442;

single ADF-462 (second optional), ALT-4000 radio altimeter, TCAS-4000 and

EGPWS Mk V. Optional add-ons include the Honeywell Laseref V IRS,

Universal Aero 1 three-channel satcom, L-3 StormScope and FDR.

As for instrumentation, the G200 incorporates Rockwell Collins EFD-4077

EFIS displays all flight and EICAS information on five 18.4 cm square screens,

dual Davtron M850A digital clocks, Flight Line 8047-10 standby altimeter,

8059-2B standby ASI, Jet AI-804CE standby AI, Precision PAI-700-04 standby

compass and Hobbs 15007 hour meter.

The Gulfstream G200 is also armed with up-to-date safety equipment

such as an Enhanced Ground Proximity Warning System (EGPWS), Traffic

Alert and Collision Avoidance System (TCAS II) and an Emergency Vision

Assurance System (EVAS).

11. Cabin

As a business jet, the Gulfstream G200 boasts large and comfortable cabin

interiors with a large baggage compartment and various amenities. The

basic cabin arrangement seats 9 passengers, although the interior can be

modified to seat 16 passengers in a corporate jet seating plan.

Figure 24. Default cabin seating arrangement

Cabin amenities include passenger service units comprising of reading and

table lights, swivel air outlets, audio system speakers and individual

headphone controls. In terms of multimedia, the cabin has a 38 cm LCD

monitor in forward cabin bulkhead, a MagnaStar digital telephone system,

three 110 V power outlets with adjacent data ports and an Airshow 410

passenger flight information system. The G200 also boasts a full aft

lavoratory equipped with a 110 V outlet pressurized water tank with a

standard capacity of 19 liters. As for the baggage compartment in the aft

fuselage, it boasts a capacity of 4.24 m3 and can be accessed via an external

airstair door.

Figure 25. Cabin interior of Gulfstream G200 in default 9 passenger arrangement

12. References

Gulfstream G200: Jane’s All The World Aircraft (Year Unknown)

Date Accessed: 14/4/2012

http://www2.janes.com.wwwproxy0.library.unsw.edu.au/janesdata/yb/jawa/

jawa5514.htm#toclink-j0010120142434

Jet Advisors: Private Jet Solutions (Year Unknown)


http://www.jetadvisors.com/aircrafts/gulfstream200galaxy.htm

Gulfstream G200: Smart Cockpit, ‘Flight Controls’


http://www.smartcockpit.com/pdf/plane/gulfstream/G200/systems/0011/

Gulfstream G200: Smart Cockpit, ‘Landing Gear’


http://www.jetadvisors.com/aircrafts/gulfstream200galaxy.htm

http://www2.janes.com.wwwproxy0.library.unsw.edu.au/janesdata/yb/jawa/jawa5514.htm#toclink-j0010120142434

http://www2.janes.com.wwwproxy0.library.unsw.edu.au/janesdata/yb/jawa/jawa5514.htm#toclink-j0010120142434



Aircraft Spruce: Everything for Planes and Pilots


http://www.aircraftspruce.com/menus/lg/tirestubes_michelin.html

Gulfstream G200: Smart Cockpit, ‘Thrust Reverser System’



Glenn Research Center: NASA (2010), “What is lift?”

Date Accessed: 10/5/2012http://www.grc.nasa.gov/WWW/K-12/airplane/lift1.html

Kroo, I. (Year Unknown), ‘High Lift Systems- Introduction’, Course Notes of

Aircraft Design course in Stanford University

Date Accessed: 10/5/2012http://adg.stanford.edu/aa241/highlift/highliftintro.html

Glenn Research Center: NASA (2010), “Flaps and Slats”

Date Accessed: 10/5/2012http://www.grc.nasa.gov/WWW/k-12/airplane/flap.html

Airframes.org (2011)

Date Accessed: 10/5/2012http://www.airframes.org/

Kroo, I. and Alonso, J. (Year Unknown) ‘Engine Placement’, Course Notes of

http://www.airframes.org/

http://www.grc.nasa.gov/WWW/k-12/airplane/flap.html

http://adg.stanford.edu/aa241/highlift/highliftintro.html

http://www.grc.nasa.gov/WWW/K-12/airplane/lift1.html


http://www.aircraftspruce.com/menus/lg/tirestubes_michelin.html


Aircraft Design course in Stanford University

Date Accessed: 12/5/2012http://adg.stanford.edu/aa241/propulsion/engineplacement.html

Pratt & Whitney Canada: PW 306A


http://www.pwc.ca/en/engines/pw306a

Gulfstream G200: Smart Cockpit, ‘Avionic Systems



Gulfstream G200: Smart Cockpit, ‘Electrical Power Systems’



Gulfstream G200: Smart Cockpit, ‘Environmental Control Systems’



Gulfstream G200: Smart Cockpit, ‘Fuel System’



Gulfstream G200: Smart Cockpit, ‘Ice & Rain Protection’








http://www.pwc.ca/en/engines/pw306a

http://adg.stanford.edu/aa241/propulsion/engineplacement.html

Documents

Gulfstream G 200