Aircraft pressurization system

Cabin pressurization provides a comfortable environment for passengers and crew while

allowing the aircraft to fly at higher altitudes. Flying at high altitudes is more fuel efficient and it

allows the aircraft to fly above most undesirable weather conditions. If an aircraft is to be

pressurized, the pressurized section (pressure vessel) must be strong enough to withstand

operational stresses. In general, the maximum altitude at which an aircraft can fly is limited by

the maximum allowable cabin differential pressure (pressure difference between ambient air and

the air inside the pressure vessel).

Turbine-engine aircraft usually utilize engine bleed air for pressurization. In these systems, high

pressure air is “bled” from the turbine-engines compressor. This also causes a reduction in

engine power but it is not as significant of a loss. Some aircraft use independent cabin

compressors for pressurization which are used to eliminate the problem of air contamination

Independent cabin compressors are driven by either:

1. The engine accessory section

2. Turbine-engine bleed air

These compressors may use one of two types of pumps:

1. Roots-type positive displacement pumps

2. Centrifugal cabin compressors

Components of pressurization system:

1. Cabin pressure regulator: The cabin pressure regulator controls cabin pressure to a

selected value in the isobaric range and limits cabin pressure to a preset differential value

in the differential range. When the airplane reaches the altitude at which the difference

between the pressure inside and outside the cabin is equal to the highest differential

pressure for which the fuselage structure is designed, a further increase in airplane altitude

will result in a corresponding increase in cabin altitude. Differential control is used to

prevent the maximum differential pressure, for which the fuselage was designed, from

being exceeded. This differential pressure is determined by the structural strength of the

cabin and often by the relationship of the cabin size to the probable areas of rupture, such

as window areas and doors.

2. Heat exchanger: It is used to cool the hot pressurized air to a usable temperature.

3. Outflow valve: The outflow valve regulates the amount of pressurized air that is allowed

to exit the cabin. Hence is provides a constant inflow of air to the pressurized area.

4. Cabin pressure safety valve: The cabin air pressure safety valve is a combination pressure

relief, vacuum relief, and dump valve. The pressure relief valve prevents cabin pressure

from exceeding a predetermined differential pressure above ambient pressure. The

vacuum relief prevents the cabin pressure from going below that of the ambient pressure

by allowing external air to enter the cabin when cabin pressure goes below the ambient

pressure. Dump valve is used to release all cabin pressure when aircraft lands. It is often

controlled by landing gear squat switch. When this switch is positioned to ram, a solenoid

valve opens, causing the valve to dump cabin air to atmosphere.

Other than the above components an aircraft pressurization system is incorporated with a cabin

differential pressure gauge which indicates the difference between inside and outside pressure.

This gauge should be monitored to assure that the cabin does not exceed the maximum allowable

differential pressure. A cabin altimeter is also provided as a check on the performance of the

system. Sometimes these two instruments combined with each other. Another instrument is also

incorporated which indicates the cabin rate of climb or descent.

In order to pressurize the cabin, air is bled from the compressor directly. Hot air is passed

through the heat exchanger where it is cool down to a temperature usable for cabin

pressurization. The cool air is then channeled to a paper filter and thence to cabin pressure

regulator. The regulator senses the pressure differential and controls the pressure inside the

cabin. Excess air is bled directly to the cabin. Numerous safety devices are built into the system

to ensure proper operation. The cabin safety valve is set to 0.34 bar. The amount of pressure

differential is governed by outflow valve. This system ensures a constant differential pressure of

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Aircraft fuel system

The fuel system for gas turbine engine delivers a uniform flow of clean fuel at proper pressure

and in the necessary quantity required for operation of engine to the engine fuel metering system.

Despite widely varying atmospheric conditions, the fuel supply must be adequate and continuous

to meet the demands of the engine during flight.

The primary requirements of the fuel system are-

To control accurate and stable engine speed for steady-state operations and to provide

transient control to achieve rapid power changes. This system pumps, filters, and meters fuel

in response to power-available spindle (PAS) position, load-demand spindle (LDS) position,

sensed engine variables, and torque motor input from the electrical control unit.

To position the compressor variable stator vanes throughout the engine operating range to

achieve the required airflow and stall margin performance by the compressor.

To provide a schedule for starting bleed valve.

To provide automatic start schedules from sea level to 20,000 feet altitude.

To protect the engine against destructive gas generator and power turbine overspeed.

These requirements must be met over the full engine operating envelope and environment.

Components of an aircraft fuel system

1. Fuel pumps : Main fuel pumps deliver a continuous supply of fuel at the proper pressure

during operation of the aircraft engine. These engine-driven fuel pumps must be able to

deliver the maximum flow needed at high pressure to obtain satisfactory nozzle spray and

accurate fuel regulation. These fuel pumps may be divided into two distinct system

categories constant displacement and variable displacement. Gear-type constant

displacement pumps have approximately straight-line flow characteristics, whereas fuel

requirements fluctuate with flight or ambient air conditions. With variable displacement

pumps the pump displacement is changed to meet varying fuel flow requirements; that is,

the amount of fuel discharged from the pump. Hence the unit controls the amount of

applicable fuel automatically and accurately regulates the pump pressure and delivery to

the engine. Other than the main fuel pump a boost pump is also incorporated as added

reliability to the fuel system.

2. Fuel filters: All gas turbine engines have several fuel filters at various points along the

system. It is common practice to install at least one filter before the fuel pump and one on

the high-pressure side of the pump. In most cases the filter will incorporate a relief valve

set to open at a specified pressure differential to provide a bypass for fuel when filter

contamination becomes excessive.

3. Pressurizing and drain(dump) valve: The pressurizing and drain valve prevents flow to

the fuel nozzles until sufficient pressure is reached in the main fuel control. Once

pressure is attained, the servo assemblies compute the fuel-flow schedules. It also drains

the fuel manifold at engine shutdown to prevent post-shutdown fires. But it keeps the

upstream portion of the sure and delivery to the engine system primed to permit faster

starts.

4. Fuel heater: Gas turbine engine fuel systems are very susceptible to the formation of ice

in the fuel filters. A fuel heater operates as a heat exchanger to warm the fuel. The heater

can use engine bleed air, an air-to-liquid exchanger, or an engine lubricating oil, a liquid-

to-liquid exchanger, as a source of heat. A fuel heater protects the engine fuel system

from ice formation.

5. Fuel nozzles: On most gas turbine engines, fuel is introduced into the combustion

chamber through a fuel nozzle. This nozzle creates a highly atomized, accurately shaped

spray of fuel for rapid mixing and combustion with the primary airstream under varying

conditions of fuel and airflow. Most engines use either the single (simplex) or the dual

(duplex) nozzle.

6. Fuel shutoff valves: A fuel shutoff valve is usually installed between the fuel control unit

and the fuel nozzles. It is controlled from the pilot's compartment. When the throttle is

placed in the closed position, this ensures positive shutoff of fuel to the engine.

Figure: Schematic of Gas turbine fuel system

Operation

The gas turbine fuel system is entirely controlled by the Turbine Digital Electronic Control

Unit. Operation of fuel system can be classified as following:

Start-up

When  the start-up function is selected on the control and indication panel, ignition

solenoid valve opens which supplies the starting injectors with fuel which is injected into

the combustion chamber through fuel nozzles. Atomized fuel mixes with the air and gets

ignited on contact with the igniters. After two seconds, the fuel shut-off solenoid valve is

energized and a reduced fuel flow passes through the injection manifold and ignites in the

combustion chamber. After five seconds, the ignition solenoid valve and the two igniters

are de-energized. From this point onwards, the fuel gas passes only through the injection

manifold.

Normal operation

Under normal operation, the control unit acts on the metering valve in order to increase or

decrease the flow of gas entering the combustion chamber and adapt the power to the set

value.

Shut down

Any shut-down order given by the pilot, or initiated by any of the safety functions, results

in the closing of the fuel metering valve and the de-energization of the fuel shut-off

solenoid valve, which also closes.

Figure: Gas turbine fuel system

Lubrication system

The lubrication system is required to provide lubrication and cooling for all gears, bearings and

splines. It must also be capable of collecting foreign matter which, if left in a bearing housing or

gearbox, can cause rapid failure. Additionally, the oil must protect the lubricated components

which are manufactured from non-corrosion resistant materials. Additional requirements for a

lubrication system for a turbo-propeller engine are lubrication of heavily loaded propeller

reduction gears and high pressure oil supply to operate propeller pitch control mechanism. All

gas turbine engines are equipped with a self-contained re-circulating lubrication system in which

the oil is distributed around the engine and returned to the oil tank by pumps. The principal

components of a gas turbine lube oil system are as follows:

1. Sump: Sump. Most gas turbines have a self-contained lube oil sump. If additional oil

capacity is required, a separately mounted tank may be supplied.

2. Cooling system: Gas turbine oil may be cooled by water, air, fuel, or oil. In aviation gas

turbines, oil is cooled by passing air over the oil sump or by an oil-to-air heat exchanger.

Regenerative (fuel) cooling can be employed on gas turbines where the cooling

requirements are relatively low.

3. Venting: To prevent excessive oil loss from venting oil vapor overboard, bearing sumps

are usually vented to an air-to-oil separator. The sump air is vented to the exhaust after

passing through the separator and the oil is returned to the main sump.

4. Lube oil filters: Filters are integrally located prior to each critical feed point. A main

lube filter assembly, usually composed of a coarse filter followed by a fine filter, is

supplied with each engine. To prevent continued use of unfiltered oil, routed through the

bypass to the engine when the filters are clogged, most filters and strainers have

differential pressure gages to assist the operator in determining when the elements require

changing, and alarms to warn operators of bypassing flow.

5. Lube oil strainers: Lube oil strainers usually contain a built-in pressure relief valve of a

size sufficient to bypass all the oil around the strainer in the event of clogging so an

uninterrupted oil flow to the engine will be maintained. The bypass line should be

connected to an audible alarm to inform engine operators that strainers are clogged.

Figure: Scavenging oil system as part of re-circulatory lubrication system

There are two basic recirculatory systems, known as the ‘pressure relief valve system’ and the

‘full flow’ system. The major difference between them is in the control of the oil flow to the

bearings.

In the pressure relief valve system the oil flow to the bearing chambers is controlled by

limiting the pressure in the feed line to a given design value. This is accomplished by the use

of a spring loaded valve which allows oil to be directly returned from the pressure pump

outlet to the oil tank, or pressure pump inlet, when the design value is exceeded. The valve

opens at a pressure which corresponds to the idling speed of the engine, thus giving a

constant feed pressure over normal engine operating speeds. However, increasing engine

speed causes the bearing chamber pressure to rise sharply. This reduces the pressure

difference between the bearing chamber and feed jet, thus decreasing the oil flow rate to the

bearings as engine speed increases. To alleviate this problem, some pressure relief valve

systems use the increasing bearing chamber pressure to augment the relief valve spring load,

This maintains a constant flow rate at the higher engine speeds by increasing the pressure in

the feed line as the bearing chamber pressure increases. Although the pressure relief valve

system operates satisfactorily for engines which have a low bearing chamber pressure, which

does not unduly increase with engine speed, it becomes an undesirable system for engines

which have high chamber pressures.

The full flow system achieves the desired oil flow rates throughout the complete engine

speed range by dispensing with the pressure relief valve and allowing the pressure pump

delivery pressure to supply directly the oil feed jets. The pressure pump size is determined by

the flow required at maximum engine speed. To prevent high oil pressures from damaging

filters or coolers, pressure limiting valves are fitted to by-pass these units. These valves

normally only operate under cold starting conditions or in the event of a blockage. Advance

warning of a blocked filter may be indicated in the cockpit by a differential pressure switch

which senses an increase in the pressure difference between the inlet and outlet of the filter.

Ignition system

Ignitions systems for gas turbine engines are required to operate for starting only. Ignition

systems for a gas turbine engine usually consist of three main components:

1. Exciter box- Exciter box sends the high voltage current to ignition lead

2. Ignition lead- Ignition lead transfers the high voltage current to igniter.

3. Igniter- Igniter is mounted on the engine in such a way that it protrudes into the

combustion chamber. High voltage current generated by the exciter is discharged across

the electrodes of the igniter and ignites the fuel air mixture inside the combustion

chamber.

Usually, gas turbine engines are equipped with two or more igniter plugs. An important

characteristic of a gas turbine engine ignition system is high energy discharge at igniter plug

because it is difficult to ignite the air-fuel mixture under some operating condition like at high

altitude. The high energy discharge is accomplished by means of a storage capacitor in what is

termed as “high-energy capacitor discharge system”.

Ignition units are rated in joules and these are designed to provide variable output according to

the operation requirements. A high value output such as 12J is necessary to ensure the reliability

of operation of ignition system at high altitude and sometimes at the time of starting.

Figure: Gas turbine ignition system

Starting system

Two separate systems are required to ensure that a gas turbine engine will start satisfactorily.

Firstly, provision must be made for the compressor and turbine to be rotated up to a speed at

which adequate air passes into the combustion system to mix with fuel from the fuel spray

nozzles. Secondly, provision must be made for ignition of the air/fuel mixture in the combustion

system. During engine starting the two systems must operate simultaneously. Starters for gas

turbines engines may be classified as air turbine (pneumatic) starters, electric starters and fuel-air

combustion starters.

Air turbines are the most common starters in large aircrafts because the high pressure air is

usually available and they have the lowest weight and size. Air turbine module comprises of inlet

valve, intake, usually a single stage axial flow turbine and an exhaust. For aircraft engines the

compressed air is supplied to the turbine from either an on-board APU, a ground start cart or by

cross bleed from another engine. The turbine cranks the engine HP (high pressure) spool via

reduction gear, clutch, engine accessory gearbox, a radial shaft and a bevel gear.

Figure: Air turbine starters

Electric starters are preferred for small gas-turbine engines. Major components of an electric

system are batteries and an electric motor driving the HP spool via a clutch, gear box and bevel

gear.

Figure: Major components of an electric starter

A combustor starter unit consists of a small combustion chamber into which high pressure air,

from an aircraft-mounted storage bottle, and fuel, from the engine fuel system, are introduced.

Control valves are provided to regulate the air supply which pressurizes a fuel accumulator to

give sufficient fuel pressure for atomization and also activates the continuous ignition system.

The fuel/air mixture is ignited in the combustion chamber and the resultant gas is directed onto

the turbine of the air starter. An electrical circuit is provided to shut off the air supply which in

turn terminates the fuel and ignition systems on completion of the starting cycle.