AIRCRAFT MAINTENANCE ENGINEERING

EXHAUST SECTION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:17:00 +0000

The hot gases are exhausted overboard through the exhaust diffuser section. Internally, this section supports the power turbine and aft portion of the powershaft. The exhaust diffuser is composed of an inner and outer housing, separated by hollow struts across the exhaust passage. The inner housing is capped by either a tailcone or a cover plate which provides a chamber for cooling the powershaft bearing. A typical exhaust diffuser section is shown in figure 1.29.

Figure 1.29. Exhaust Diffuser Section.

Turboshaft engines used in helicopters do not develop thrust by use of the exhaust duct. If thrust were developed by the engine exhaust gas, it would be impossible to maintain a stationary hover; therefore, helicopters use divergent ducts. These ducts reduce gas velocity and dissipate any thrust remaining in the exhaust gases. On fixed wing aircraft, the exhaust duct may be the convergent type, which accelerates the remaining gases to produce thrust which adds additional shaft horsepower to the engine rating. The combined thrust and shaft horsepower is called equivalent shaft horsepower (ESHP).

Figure 1.30. Divergent Exhaust Duct.

TURBINE SECTION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:16:00 +0000

A portion of the kinetic energy of the expanding gases is extracted by the turbine section, and this energy is transformed into shaft horsepower which is used to drive the compressor and accessories. In turboprop and turboshaft engines, additional turbine rotors are designed to extract all of the energy possible from the remaining gases to drive a powershaft.

Types of turbines. Gas turbine manufacturers have concentrated on the axial-flow turbine shown in figure 1.21. This turbine is used in all gas-turbine-powered aircraft in the Army today. However, some manufacturers are building engines with a radial inflow turbine, illustrated in figure 1.22. The radial inflow turbine has the advantage of ruggedness and simplicity, and it is relatively inexpensive and easy to manufacture when compared to the axial-flow turbine. The radial flow turbine is similar in design and construction to the centrifugal-flow compressor described in paragraph 1.19a. Radial turbine wheels used for small engines are well suited for a higher range of specific speeds and work at relatively high efficiency.

Figure 1.21. Axial-flow Turbine Rotor.

Figure 1.22. Radial Inflow Turbine.

The axial-flow turbine consists of two main elements, a set of stationary vanes followed by a turbine rotor. Axial-flow turbines may be of the single-rotor or multiple-rotor type. A stage consists of two main components: a turbine nozzle and a turbine rotor or wheel, as shown in figure 1.21. Turbine blades are of two basic types, the impulse and the reaction. Modern aircraft gas turbines use blades that have both impulse and reaction sections, as shown in figure 1.23.

Figure 1.23. Impulse-Reaction Turbine Blade.

The stationary part of the turbine assembly consists of a row of contoured vanes set at a predetermined angle to form a series of small nozzles which direct the gases onto the blades of the turbine rotor. For this reason, the stationary vane assembly is usually called the turbine nozzle, and the vanes are called nozzle guide vanes.

Single-rotor turbine. Some gas turbine engines use a single-rotor turbine, with the power developed by one rotor. This arrangement is used on engines where low weight and compactness are necessary. A single-rotor, single-stage turbine engine is shown in figure 1.24, and a multiple-rotor, multiple-stage turbine engine is shown in figure 1.25.

Figure 1.24. Single-rotor,Single-stage Turbine.

Figure 1.25. Multiple-rotor,Multiple-stage Turbine.

Multiple-rotor turbine. In the multiple-rotor turbine the power is developed by two or more rotors. As a general rule, multiple-rotor turbines increase the total power generated in a unit of small diameter. Generally the turbines used in Army aircraft engines have multiple rotors. Figure 1.26 illustrates a multistage, multiple-rotor turbine assembly.

Figure 1.26. Multirotor - Multistage Turbine.

TURBINE COMBUSTION SECTION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:15:00 +0000

Today, three basic combustion chambers are in use. They are the annular combustion chamber, the can type, and the combination of the two called the can-annular. Variations of these basic systems are used in a number of engines. The three systems are discussed individually in the following subparagraphs. The most commonly used gas turbine engine in Army aircraft is the annular reverse-Row type. The combustion section contains the combustion chambers, igniter plugs, and fuel nozzles or vaporizing tubes. It is designed to burn a fuel-air mixture and deliver the combusted gases to the turbine at a temperature which will not exceed the allowable limit at the turbine inlet.
Fuel is introduced at the front end of the burner in a highly atomized spray from the fuel nozzles. Combustion air flows in around the fuel nozzle and mixes with the fuel to form a correct fuel-air mixture. This is called primary air and represents approximately 25 percent of total air taken into the engine. The fuel-air mixture which is to be burned is a ratio of 15 parts of air to 1 part of fuel by weight. The remaining 75 percent of the air is used to form an air blanket around the burning gases and to lower the temperature. This temperature may reach as high as 3500° F. By using 75 percent of the air for cooling, the temperature operating range can be brought down to about half, so the turbine section will not be destroyed by excessive heat. The air used for burning is called primary air- and that for cooling is secondary air. The secondary air is controlled and directed by holes and louvers in the combustion chamber liner.
Igniter plugs function only during starting, being cut out of the circuit as soon as combustion is self-supporting. On engine shutdown, or, if the engine fails to start, the combustion chamber drain valve, a pressure-actuated valve, automatically drains any remaining unburned fuel from the combustion chamber. All combustion chambers contain the same basic elements: a casing or outer shell, a perforated inner liner or flame tube, fuel nozzles, and some means of initial ignition. The combustion chamber must be of light construction and is designed to burn fuel completely in a high velocity airstream. The combustion chamber liner is an extremely critical engine part because of the high temperatures of the flame. The liner is usually constructed of welded high-nickel steel. The most severe operating periods in combustion chambers are encountered in the engine idling and maximum rpm ranges. Sustained operation under these conditions must be avoided to prevent combustion chamber liner failure.

The annular-type combustion chamber shown in figure 1.18 is used in engines of the axial-centrifugal-flow compressor design. The annular combustion chamber permits building an engine of a small and compact design. Instead of individual combustion chambers, the primary compressed air is introduced into an annular space formed by a chamber liner around the turbine assembly. A space is left between the outer liner wall and the combustion chamber housing to permit the flow of secondary cooling air from the compressor. Primary air is mixed with the fuel for combustion. Secondary (cooling) air reduces the temperature of the hot gases entering the turbine to the proper level by forming a blanket of cool air around these hot gases.

1. ANNULAR TYPE COMBUSTION CHAMBER LINER
2. COMBUSTION CHAMBER HOUSING ASSEMBLY
Figure 1.18. Annular-type Combustion Chamber.

The annular combustion chamber offers the advantages of a larger combustion volume per unit of exposed area and material weight, a smaller exposed area resulting in lower pressure losses through the unit, and less weight and complete pressure equalization.

The can-type combustion chamber is one made up of individual combustion chambers. This type of combustion chamber is so arranged that air from the compressor enters each individual chamber through the adapter. Each individual chamber is composed of two cylindrical tubes, the combustion chamber liner and the outer combustion chamber, shown in figure 1.19. Combustion takes place within the liner. Airflow into the combustion area is controlled by small louvers located in the inner dome, and by round holes and elongated louvers along the length of the liner. Airflow into the combustion area is controlled by small louvers located in the inner dome, and by round holes elongated louvers along the length of the liner.

Figure 1.19. Can-type Combustion Chamber (Cutaway).

Through these openings flows the air that is used in combustion and cooling. This air also prevents carbon deposits from forming on the inside of the liner. This is important, because carbon deposits can block critical air passages and disrupt airflow along the liner walls causing high metal temperatures and short burner life.
Ignition is accomplished during the starting cycle. The igniter plug is located in the combustion liner adjacent to the start fuel nozzle. The Army can-type engine employs a single can-type combustor.

Can-annular combustion chamber. This combustion chamber uses characteristics of both annular and can-type combustion chambers. The can-annular combustion chamber consists of an outer shell, with a number of individual cylindrical liners mounted about the engine axis as shown in figure 1.20. The combustion chambers are completely surrounded by the airflow that enters the liners through various holes and louvers. This air is mixed with fuel which has been sprayed under pressure from the fuel nozzles. The fuel-air mixture is ignited by igniter plugs, and the flame is then carried through the crossover tubes to the remaining liners. The inner casing assembly is both a support and a heat shield; also, oil lines run through it.

Figure 1.20. Can-Annular Combustion Chamber.

COMPRESSOR SECTION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:14:00 +0000

The compressor is the section of the engine that produces an increase in air pressure. It is made up of rotating and stationary vane assemblies. The first stage compressor rotor blades accelerate the air rearward into the first stage vane assemblies. The first stage vane assemblies slow the air down and direct it into the second stage compressor rotor blades. The second stage compressor rotor blades accelerate the air rearward into the second stage vane assemblies, and so on through the compressor rotor blades and vanes until air enters the diffuser section. The highest total air velocity is at the inlet of the diffuser. As the air passes rearward through the diffuser, the velocity of the air decreases and the static pressure increases. The highest static pressure is at the diffuser outlet.
The compressor rotor may be thought of as an air pump. The volume of air pumped by the compressor rotor is basically proportional to the rotor rpm. However, air density, the weight of a given volume of air, also varies this proportional relationship. The weight per unit volume of air is affected by temperature, compressor air inlet pressure, humidity, and ram air pressure*. If compressor air inlet temperature is increased, air density is reduced. If compressor air inlet pressure is increased, air density is increased. If humidity increases, air density is decreased. Humidity, by comparison with temperature, and pressure changes, has a very small effect on density. With increased forward speed, ram air pressure increases and air temperature and pressure increase.
*ram air pressure - free stream air pressure provided by the forward motion of the engine.
The following is an example of how air density affects compressor efficiency of the T62 gas turbine. At 100 percent N1 rpm, the compressor rotor pumps approximately 40.9 cubic feet of air per second. At standard day static sea level conditions, 59° F outside air temperature and 29.92" Hg barometric pressure, with 0 percent relative humidity and 0 ram air, air density is .07651 pound per cubic foot. Under these conditions, 40.9 cubic feet per second times .07651 pound per cubic feet equals approximately 3.13 pounds per second air flow through the engine. If the air density at the compressor inlet is less than on a standard day, the weight of air flow per second through the engine is less than 3.13 pounds per second. If N1 is less than 100 percent rpm on a standard day, the weight of air flow per second through the engine will be less than 3. 13 due to decreased volume flow at lower rpm. Because of this, N1 rpm varies with the power output. If output power is increased, N1 rpm will increase and vice versa. Thus, the weight of air pumped by the compressor rotor is determined by rpm and air density.
Compressor efficiency determines the power necessary to create the pressure rise of a given airflow, and it affects the temperature change which takes place in the combustion chamber. Therefore, the compressor is one of the most important components of the gas turbine engine because its efficient operation is the key to overall engine performance. The following subparagraphs discuss the three basic compressors used in gas turbine engines: the centrifugal-flow, the axial-flow, and axial-centrifugal-flow compressors. The axial-centrifugal-flow compressor is a combination of the other two and operates with characteristics of both.

Centrifugal-flow compressor. Figure 1.12 shows the basic components of a centrifugal-flow compressor: rotor, stator, and compressor manifold.

Figure 1.12. Typical Single-stage Centrifugal Compressor

As the impeller (rotor) revolves at high speed, air is drawn into the blades near the center. Centrifugal force accelerates this air and causes it to move outward from the axis of rotation toward the rim of the rotor where it is forced through the diffuser section at high velocity and high kinetic energy. The pressure rise is produced by reducing the velocity of the air in the diffuser, thereby converting velocity energy to pressure energy. The centrifugal compressor is capable of a relatively high compression ratio per stage. This compressor is not used on larger engines because of size and weight.
Because of the high tip speed problem in this design, the centrifugal compressor finds its greatest use on the smaller engines where simplicity, flexibility of operation, and ruggedness are the principal requirements rather than small frontal area and ability to handle high airflows and pressures with low loss of efficiency.

Axial-flow compressor. The air is compressed, as the name implies, in a direction parallel to the axis of the engine. The compressor is made of a series of rotating airfoils called rotor blades, and a stationary set of airfoils called stator vanes. A stage consists of two rows of blades, one rotating and one stationary. The entire compressor is made up of a series of alternating rotor and stator vane stages as shown in figure 1.13.

Figure 1.13. Axial-flow Compressor.

Axial flow compressors have the advantage of being capable of very high compression ratios with relatively high efficiencies; see figure 1.14. Because of the small frontal area created by this type of compressor, it is ideal for installation on high-speed aircraft. Unfortunately, the delicate blading and close tolerances, especially toward the rear of the compressor where the blades are smaller and more numerous per stage, make this compressor highly susceptible to foreign-object damage. Because of the close fits required for efficient air-pumping and higher compression ratios, this type of compressor is very complex and very expensive to manufacture. For these reasons the axial-flow design finds its greatest application where required efficiency and output override the considerations of cost, simplicity, and flexibility of operation. However, due to modern technology, the cost of the small axial-flow compressor, used in Army aircraft, is coming down.

Figure 1.14. Compressor Efficiencies and Pressure Ratios.

Axial-centrifugal-flow compressor. The axial-centrifugal-flow compressor, also called the dual compressor, is a combination of the two types, using the same operating characteristics. Figure 1.15 shows the compressor used in the T53 turbine engine. Most of the gas turbine engines used in Army aircraft are of the dual compressor design. Usually it consists of a five- or seven-stage axial-flow compressor and one centrifugal-flow compressor. The dual compressors are mounted on the same shaft and turn in the same direction and at the same speed. The centrifugal compressor is mounted aft of the axial compressor. The axial compressor contains numerous air-foil-shaped blades and vanes that accomplish the task of moving the air mass into the combustor at an elevated pressure.

Figure 1.15. Axial-Centrifugal-Flow Compressor.

As the air is drawn into the engine, its direction of flow is changed by the inlet guide vanes. The angle of entry is established to ensure that the air flow onto the rotating compressor blades is within the stall-free (angle of attack) range. Air pressure or velocity is not changed as a result of this action. As the air passes from the trailing edge of the inlet guide vanes, its direction of flow is changed due to the rotational effect of the compressor. This change of airflow direction is similar to the action that takes place when a car is driven during a rain or snow storm. The rain or snow falling in a vertical direction strikes the windshield at an angle due to the horizontal velocity of the car.
In conjunction with the change of airflow direction, the velocity of the air is increased. Passing through the rotating compressor blades, the velocity is decreased, and a gain in pressure is obtained. When leaving the trailing edge of the compressor blades, the velocity of the air mass is again increased by the rotational effect of the compressor. The angle of entry on to the stationary stator vanes results from this rotational effect as it did on the airflow onto the compressor.
Passing through the stationary stator vanes the air velocity is again decreased resulting in an increase in pressure. The combined action of the rotor blades and stator vanes results in an increase in air -pressure; combined they constitute one stage of compression. This action continues through all stages of the axial compressor. To retain this pressure buildup, the airflow is delivered, stage by stage, into a continually narrowing airflow path. After passing from the last set of stator vanes the air mass passes through exit guide vanes. These vanes direct the air onto the centrifugal impeller.

The centrifugal impeller increases the velocity of the air mass as it moves it in a radial direction.

Compressor stall. Gas turbine engines are designed to avoid the pressure conditions that allow engine surge to develop, but the possibility of surge still exists in engines that are improperly adjusted or have been abused. Engine surge occurs any time the combustion chamber pressure exceeds that in the diffuser, and it is identified by a popping sound which is issued from the inlet. Because there is more than one cause for surge, the resultant sound can range from a single carburetor backfire pop to a machinegun sound.

Engine surge is caused by a stall on the airfoil surfaces of the rotating blades or stationary vanes of the compressor. The stall can occur on individual blades or vanes or, simultaneously, on groups of them. To understand how this can induce engine surge, the causes and effects of stall on any airfoil must be examined.
All airfoils are designed to provide lift by producing a lower pressure on the convex (suction) side of the airfoil than on the concave (pressure) side. A characteristic of any airfoil is that lift increases with an increasing angle of attack, but only up to a critical angle. Beyond this critical angle of attack, lift falls off rapidly. This is due largely to the separation of the airflow from the suction surface of the airfoil, as shown in the sketch. This phenomenon is known as stall. All pilots are familiar with this condition and its consequences as it applies to the wing of an aircraft. The stall that takes place on the fixed or rotating blades of a compressor is the same as the stalling phenomenon of an aircraft wing.

TURBINE AIR INLET SECTION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:13:00 +0000

The amount of air required by a gas turbine engine is approximately ten times that of a reciprocating engine. The air inlet is generally a large, smooth aluminum or magnesium duct which must be designed to conduct the air into the compressor with minimum turbulence and restriction. The air inlet section may have a variety of names according to the desire of the manufacturer. It may be called the front frame and accessory section, the air inlet assembly, the front bearing support and shroud assembly, or any other term descriptive of its function. Usually, the outer shell of the front frame is joined to the center portion by braces that are often called struts. The anti-icing system directs compressor discharge air into these struts. The temperature of this air prevents the formation of ice that might prove damaging to the engine. Anti-icing systems are discussed further in the lesson covering the engines they may be installed on.

ADVANTAGES OF TURBINE ENGINES

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:07:00 +0000

Keeping in mind the basic theory of turbine engines, compare the advantages and disadvantages of the turbine engine with the piston or reciprocating engine. The advantages are covered in the subparagraphs below, and disadvantages are discussed in the next section.

Power-to-weight ratio. Turbine engines have a higher power-to-weight ratio than reciprocating engines. An example of this is the T55-L-l11. It weighs approximately 650 pounds and delivers 3, 750 shaft horsepower. The power-to-weight ratio for this engine is 5.60 shp per pound, where the average reciprocating engine has a power-to-weight ratio of approximately .67 shp per pound.
Less maintenance. Maintenance per hour of operation is especially important in military operations. Turbine engines require less maintenance per flying hour than reciprocating engines generally do. As an aircraft maintenance officer, this advantage will appeal to you because of a greater aircraft availability and lower maintenance hour to flying hour ratio. The turbine engine also has fewer moving parts than a reciprocating engine; this is also an advantage over the reciprocating engine.
Less drag. Because of the design, the turbine engine has a smaller frontal area than the reciprocating engine. A reciprocating engine requires a large frontal area which causes a great deal of drag on the aircraft. Turbine engines are more streamlined in design, causing less drag. Figure 1.6 shows one of the two nacelles that contain reciprocating engines in the old CH-37 cargo helicopter. Figure 1.7 shows the smaller frontal area of the turbine engines that power the CH-47 Chinook helicopter. Because of this, the engine nacelles are more streamlined in design, causing less drag.
Cold weather starting. The turbine engine does not require any oil dilution or preheating of the engine before starting. Also, once started, the reciprocating engine takes a long time to warm up to operating temperatures, whereas the turbine engine starts readily and is up to operating temperature immediately.
Low oil consumption. The turbine engine, in general, has a lower rate of oil consumption than the reciprocating engine. The turbine engine does not require the oil reservoir capacity to be as large as the reciprocating engine's; because of this, a weight and economy factor is an additional advantage.

TURBOJET

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:05:00 +0000

The turbojet is the engine in most common use today in high-speed, high-altitude aircraft, not in Army aircraft. With this engine, air is drawn in by a compressor which raises internal pressures many times over atmospheric pressure. The compressed air then passes into a combustion chamber where it is mixed with fuel to be ignited and burned. Burning the fuel-air mixture expands the gas, which is accelerated out the rear as a high-velocity jet-stream. In the turbine section of the engine, the hot expanded gas rotates a turbine wheel which furnishes power to keep the compressor going. The gas turbine engine operates on the principle of intake, compression, power, and exhaust, but unlike the reciprocating engine, these events are continuous. Approximately two-thirds of the total energy developed within the combustion chamber is absorbed by the turbine wheel to sustain operation of the compressor. The remaining energy is discharged from the rear of the engine as a high velocity jet, the reaction to which is thrust or forward movement of the engine. The turbojet is shown schematically in figure 1.3.

Figure 1.3. Axial-Flow Turbojet Engine.

BRAYTON CYCLE OF OPERATION

msmprr@yahoo.com (Aviation) — Sat, 30 Jul 2011 12:03:00 +0000

Ambient air is drawn into the inlet section by the rotating compressor. The compressor forces this incoming air rearward and delivers it to the combustion chamber at a higher pressure than the air had at the inlet. The compressed air is then mixed with fuel that is sprayed into the combustion chamber by the fuel nozzles. The fuel and air mixture is then ignited by electrical igniter plugs similar to spark plugs. This ignition system is only in operation during the starting sequence, and once started, combustion is continuous and self-sustaining as long as the engine is supplied with the proper air-fuel ratio. Only about 25 percent of the air is used for combustion. The remaining air is used for internal cooling and pressurizing.
The turbine engines in the Army inventory are of the free-power turbine design, as shown in figure 1.2. In this engine, nearly two-thirds of the energy produced by combustion is extracted by the gas producer turbine to drive the compressor rotor. The power turbine extracts the remaining energy and converts it to shaft horsepower (shp), which is used to drive the output shaft of the engine. The gas then exits the engine through the exhaust section to the atmosphere. Army helicopters use a divergent duct to eliminate the remaining thrust. The various kinds of exhaust ducting are discussed in detail with the engine using that particular ducting.

Figure 1.2. Typical Free-Power Turboshaft Engine

Thermal expansion valve

msmprr@yahoo.com (Aviation) — Fri, 10 Jun 2011 05:59:00 +0000

Thermal expansion valve(ATA 21)

A thermal expansion valve is a component in refrigeration and air conditioning systems that controls the amount of refrigerant flow into the evaporator thereby controlling the superheat at the outlet of the evaporator. This is accomplished by use of a temperature sensing bulb filled with a similar gas as in the system that causes the valve to open against the spring pressure in the valve body as the temperature on the bulb increases. As temperatures in the evaporator decrease, so does the pressure in the bulb and therefore on the spring causing the valve to close. An air conditioning system with a TX valve is often more efficient than other designs that do not use one.

Self-sealing fuel tank

msmprr@yahoo.com (Aviation) — Wed, 23 Feb 2011 09:44:00 +0000

In aviation, self-sealing fuel tank is a fuel tank technology in wide use since World War II that prevents fuel tanks primarily on aircraft from leaking fuel and igniting after being damaged by enemy fire.

Self-sealing tanks have two layers of rubber, one of vulcanized rubber and one of untreated rubber that can absorb oil and expand when wet. When a fuel tank is punctured, the fuel will spill on to the layers, causing the swelling of the untreated layer, thus sealing the puncture.

Most jet fighters and all US military rotary wing aircraft have some type of self-sealing tanks. Military rotary wing fuel tanks have the additional feature of being crashworthy. High altitudes require the tanks to be pressurized, making self-sealing difficult. Newer technologies have brought advances like inert foam-filled tanks to prevent detonation. This foam is an open cell foam that effectively divides the gas space above the remaining fuel into thousands of small spaces; none of which contain sufficient vapour to support combustion. This foam also serves to reduce fuel slosh. Major manufacturers of this technology include Amfuel (Zodiac) (formerly Firestone), Engineered Fabrics Corp. (Meggitt) (formerly Goodyear), GKN USA and FPT Industries. For military use, tanks are qualified to MIL-DTL-27422 and MIL-DTL-5578.

In additions to fighter aircraft some military patrol vehicles and armoured limousines for VIP use also feature self-sealing fuel tanks.

Wingtip device

msmprr@yahoo.com (Aviation) — Mon, 14 Feb 2011 13:47:00 +0000

Wingtip devices are usually intended to improve the efficiency of fixed-wing aircraft. There are several types of wingtip devices, and though they function in different manners, the intended effect is always to reduce the aircraft's drag by altering the airflow near the wingtips. Wingtip devices can also improve aircraft handling characteristics and enhance safety for following aircraft. Such devices increase the effective aspect ratio of a wing without materially increasing the wingspan. An extension of span would lower lift-induced drag, but would increase parasitic drag and would require boosting the strength and weight of the wing. At some point, there is no net benefit from further increased span. There may also be operational considerations that limit the allowable wingspan (e.g., available width at airport gates).

Wingtip devices increase the lift generated at the wingtip (by smoothing the airflow across the upper wing near the tip) and reduce the lift-induced drag caused by wingtip vortices, improving lift-to-drag ratio. This increases fuel efficiency in powered aircraft and increases cross-country speed in gliders, in both cases increasing range.

Winglets are employed on many aircraft types, such as:

Rutan VariEze, the first aircraft to use winglets (1975)
Learjet 28/29, the first production jet aircraft to use winglets (1977)
Glaser-Dirks DG-303, an early glider derivative design, incorporating winglets as factory standard equipment
Airbus A310-300, the first airliner to feature wingtip fences (1985)
Boeing 747-400, the first mainline airliner to feature winglets (1988)
Ilyushin Il-96, first Russian and modern jet to feature winglets (1988)
Tupolev Tu-204, first narrow body aircraft to feature winglets (1994)

Lightning detection

msmprr@yahoo.com (Aviation) — Tue, 28 Dec 2010 06:30:00 +0000

A lightning detector is a device that detects lightning produced by thunderstorms. There are three primary types of detectors: ground-based systems using multiple antennas, mobile systems using a direction and a sense antenna in the same location (often aboard an aircraft), and space-based systems.

For more http://avionics0.blogspot.com/2010/12/lightning-detection.html

Compare Airbus 320 family

msmprr@yahoo.com (Aviation) — Sun, 26 Dec 2010 15:05:00 +0000

Flaps

msmprr@yahoo.com (Aviation) — Sun, 14 Nov 2010 06:12:00 +0000

Flaps are hinged surfaces on the trailing edge of the wings of a fixed-wing aircraft. As flaps are extended, the stalling speed of the aircraft is reduced, which means that the aircraft can fly safely at slower speeds (especially during take off and landing). Flaps are also used on the leading edge of the wings of some high-speed jet aircraft, where they may be called Krueger flaps.
Extending flaps increases the camber of the wing airfoil, thus raising the maximum lift coefficient. This increase in maximum lift coefficient allows the aircraft to generate a given amount of lift with a slower speed. Therefore, extending the flaps reduces the stalling speed of the aircraft
Extending flaps also increases drag. This can be beneficial in the approach and landing phase because it helps to slow the aircraft. Another useful side effect of flap deployment is a decrease in aircraft pitch angle. This provides the pilot with a greater view over the nose of the aircraft and allows a better view of the runway during approach and landing.
Some trailing edge flap systems increase the planform area of the wing in addition to changing the camber. In turn, the larger lifting surface allows the aircraft to generate a given amount of lift with a slower speed, thus further reducing stalling speed. Although this effect is very similar to increasing the lift coefficient, raising the planform area of the wing does not itself raise the lift coefficient. The Fowler flap is an example of a flap system that increases the planform area of the wing in addition to increasing the camber.

Triple-slotted trailing-edge flaps and leading edge Krueger (unslotted and slotted) flaps fully extended on a Boeing 747 for landing.

Types of flap systems include:

Krueger flap: hinged flap on the leading edge. Often called a "droop".
Plain flap: rotates on a simple hinge.
Split flap: upper and lower surfaces are separate, the lower surface operates like a plain flap, but the upper surface stays immobile or moves only slightly.
Gouge flap: a cylindrical or conical aerofoil section which rotates backwards and downwards about an imaginary axis below the wing, increasing wing area and chord without affecting trim. Invented by Arthur Gouge for Short Brothers in 1936.
Fowler flap: slides backwards before hinging downwards, thereby increasing both camber and chord, creating a larger wing surface better tuned for lower speeds. It also provides some slot effect. The Fowler flap was invented by Harlan D. Fowler.
Fairey-Youngman flap: moves body down before moving aft and rotating.
Slotted flap: a slot (or gap) between the flap and the wing enables high pressure air from below the wing to re-energize the boundary layer over the flap. This helps the airflow to stay attached to the flap, delaying the stall.
Blown flaps: systems that blow engine air over the upper surface of the flap at certain angles to improve lift characteristics.

Primary Flight Display (PFD)

msmprr@yahoo.com (Aviation) — Fri, 12 Nov 2010 11:56:00 +0000

A primary flight display or PFD is a modern aircraft instrument dedicated to flight information. Much like multi-function displays, primary flight displays are built around an LCD or CRT display device. Representations of older six pack or "steam gauge" instruments are combined on one compact display, simplifying pilot workflow and streamlining cockpit layouts.

For more http://avionics0.blogspot.com/2010/11/primary-flight-display-pfd.html

Thermal Anti-Icing System

msmprr@yahoo.com (Aviation) — Thu, 21 Oct 2010 07:58:00 +0000

Thermal systems used for the purpose of preventing the formation of ice or for deicing airfoil leading edges, usually use heated air ducted spanwise along the inside of the leading edge of the airfoil and distributed around its inner surface. However, electrically heated elements are also used for anti-icing and deicing airfoil leading edges.
There are several methods used to provide heated air. These include bleeding hot air from the turbine
compressor, engine exhaust heat exchangers, and ram air heated by a combustion heater.

Anti-Icing Using Cotibustion Heaters
Anti-icing systems using combustion heaters usually have a separate system for each wing and the empennage. A typical system of this type has the required number of combustion heaters located in each wing and in the empennage. A system of ducting and valves controls the airflow. The anti-icing system is automatically controlled by overheat switches, thermal cycling switches, a balance control, and a duct pressure safety switch The overheat and cycling switches allow the heaters to operate at periodic intervals, and they also stop
heater operation completely if overheating occurs.

Anti-Icing Using Exhaust Heaters
Anti-icing of the wing and tail leading edges is accomplished by a controlled flow of heated air from heat muffs around a reciprocating engine’s tail pipe. In some installations this assembly is called an augmentor.

Normally, heated air from either engine supplies the wing leading edge anti-icing system in the same wing section. During single engine operation, a crossover duct system interconnects the left and right wing leading edge ducts. This duct supplies heated air to the wing section normally supplied by the inoperative engine. Check valves in the crossover duct prevent the reverse flow of heated air and also prevent cold air from entering the anti-icing system from the inoperative engine.

Anti Icing using Engine Bleed Air

Heated air for anti.icing is obtained by bleeding air from the engine compressor. The reason for the use of such a system is that relatively large amounts of very hot air can be tapped off the compressor, providing a satisfactory source of anti-icing and deicing heat.

The shut off valve for each anti-icing section is a pressure regulating type. The valve controls the flow of air from the bleed air system to the ejectors, where it is ejected through small nozzles into mixing chambers. The hot bleed air is mixed with ambient air.

Cabin Pressurization

msmprr@yahoo.com (Aviation) — Mon, 11 Oct 2010 11:51:00 +0000

Cabin pressurization is the active pumping of compressed air into an aircraft cabin when flying at altitude to maintain a safe and comfortable environment for crew and passengers in the low outside atmospheric pressure.
Pressurization is essential over 3,000 metres (9,800 ft) above sea level to protect crew and passengers from the risk of hypoxia and a number of other physiological problems in the thin air above that altitude and increases passenger comfort generally. "The outflow valve is constantly being positioned to maintain cabin pressure as close to sea level as practical, without exceeding a cabin-to-outside pressure differential of 8.60 psi."
Maintaining the cabin pressure altitude to below 3,000 metres (9,800 ft) generally avoids significant hypoxia, altitude sickness, decompression sickness and barotrauma. Emergency oxygen systems are installed, both for passengers and cockpit crew, to prevent loss of consciousness in the event that cabin pressure rapidly rises above 10,000 feet MSL. Those systems contain more than enough oxygen for all on board, to give the pilot adequate time to descend the plane to a safe altitude, where supplemental oxygen is not needed. Federal Aviation Administration (FAA) regulations in the U.S. mandate that the cabin altitude may not exceed 8,000 feet at the maximum operating altitude of the airplane under normal operating conditions.
The pressure maintained within the cabin is referred to as the equivalent effective cabin altitude or more normally, the "cabin altitude". Cabin altitude is not normally maintained at average mean sea level (MSL) pressure (1013.25 hPa, or 29.921 inches of mercury) throughout the flight, because doing so would cause the designed differential pressure limits of the fuselage to be exceeded. An aircraft planning to cruise at 40,000 ft (12,000 m) is programmed to rise gradually from take-off to around 8,000 ft (2,400 m) in cabin pressure altitude, and to then reduce gently to match the ambient air pressure of the destination.
Pressurization is achieved by the design of an airtight fuselage engineered to be pressurized with a source of compressed air and controlled by an environmental control system (ECS). The most common source of compressed air for pressurization is bleed air extracted from the compressor stage of a gas turbine engine, from a "low" or "intermediate" stage and also from an additional "high" stage. "The exact stage can vary, depending on engine type." By the time the cold outside air has reached the bleed air valves, it has been heated to around 200 °C (392 °F) and is at a very high pressure. The control and selection of high or low bleed sources is fully automatic and is governed by the needs of various pneumatic systems at various stages of flight.
The part of the bleed air that is directed to the ECS, is then expanded and cooled to a suitable temperature by passing it through a heat exchanger and air cycle machine ('the packs system'). In some of the larger airliners, hot trim air can be added downstream of air conditioned air coming from the packs, if it is needed to warm a section of the cabin that is colder than other sections
All exhaust air is dumped to atmosphere via an outflow valve, usually at the rear of the fuselage. This valve controls the cabin pressure and also acts as a safety relief valve, in addition to other safety relief valves. In the event that the automatic pressure controllers fail, the pilot can manually control the cabin pressure valve, according to the backup emergency procedure checklist. The automatic controller normally maintains the proper cabin pressure altitude by constantly adjusting the outflow valve position, so that the cabin pressure is as near to sea level pressure as practical, without exceeding the maximum differential limit of 8.60 psi. At 39,000 feet, the cabin pressure would be automatically maintained at about 6,900 feet (450 feet lower than Mexico City), which is about 11.5 psi of atmosphere pressure (76 kPa)

Environmental Control System (ECS)

msmprr@yahoo.com (Aviation) — Sat, 25 Sep 2010 09:29:00 +0000

The Environmental Control System of an airliner provides air supply, thermal control and cabin pressurization for the passengers and crew. Avionics cooling, smoke detection, and fire suppressionare also commonly considered part of the Environmental Control System.

On most jetliners, air is supplied to the ECS by being "bled" from a compressor stage of each gas turbine engine, upstream of the combustor. The temperature and pressure of this "bleed air" varies widely depending upon which compressor stage and the RPM of the engine.

A "Manifold Pressure Regulating Shut-Off Valve" (MPRSOV) restricts the flow as necessary to maintain the desired pressure for downstream systems. This flow restriction results in efficiency losses. To reduce the amount of restriction required, and thereby increase efficiency, air is commonly drawn from two bleed ports (3 on the Boeing 777).

When the engine is at low thrust, the air is drawn from the "High Pressure Bleed Port." As thrust is increased, the pressure from this port rises until "crossover," where the "High Pressure Shut-Off Valve" (HPSOV) closes and air is thereafter drawn from the "Low Pressure Bleed Port."

To achieve the desired temperature, the bleed-air is passed through a heat exchanger called a "pre-cooler." Air from the jet engine fan is blown across the pre-cooler, which is located in the engine strut. A "Fan Air Modulating Valve" (FAMV) varies the cooling airflow, and thereby controls the final air temperature of the bleed air.

The Cold Air Unit, or "Airconditioning pack" is usually an air cycle machine (ACM) cooling device. Some aircraft, including early 707 jetliners, used vapor-compression refrigeration like that used in home air conditioners.

On most jetliners, the A/C packs are located in the "Wing to Body Fairing" between the two wings beneath the fuselage. On some jetliners (Douglas Aircraft DC-9 Series) the A/C Packs are located in the tail. The A/C Packs on the McDonnell Douglas DC-10/MD-11 and LockheedL-1011 are located in the front of the aircraft beneath the flight deck. Nearly all jetliners have two packs, although larger aircraft such as the Boeing 747, Lockheed L-1011, and McDonnell-Douglas DC-10/MD-11 have three.

The quantity of bleed air flowing to the A/C Pack is regulated by the "Flow Control Valve" (FCV). One FCV is installed for each pack. A normally closed "isolation valve" prevents air from the left bleed system from reaching the right pack (and v.v.), although this valve may be opened in the event of loss of one bleed system.

The A/C Pack exhaust air is ducted into the pressurized fuselage, where it is mixed with filtered air from the recirculation fans, and fed into the "mix manifold". On nearly all modern jetliners, the airflow is approximately 50% "outside air" and 50% "filtered air."
Modern jetliners use "High Efficiency Particulate Arresting" HEPA filters, which trap >99% of all bacteria and clustered viruses.
Airflow into the fuselage is approximately constant, and pressure is maintained by varying the opening of the "Out Flow Valve" (OFV). Most modern jetliners have a single OFV located near the bottom aft end of the fuselage, although some larger aircraft like the 747 and 777 have two.
In the event the OFV should fail closed, at least two Positive Pressure Relief Valves (PPRV) and at least one Negative Pressure Relief Valve (NPRV) are provided to protect the fuselage from over- and under- pressurization.
The atmosphere at typical jetliner cruising altitudes is generally very dry and cold, and the outside air pumped into the cabin on a long flight typically has a relative humidity around 10%. The fact that cabin pressure is generally lower than the pressure at ground level does not of itself contribute to the dryness.

AIRCRAFT MASS AND BALANCE

msmprr@yahoo.com (Aviation) — Fri, 17 Sep 2010 04:01:00 +0000

AIRCRAFT MASS AND BALANCE

The main purposes, of monitoring the mass and balance of aircraft, are to maintain safety and to achieve efficiency in flight. The position of loads such as passengers, fuel, cargo and equipment will alter the position of the Centre of Gravity (CG) of the aircraft.

Incorrect loading will affect the aircraft rate of climb, manoeuvrability, ceiling, speed and fuel consumption. If the CG were too far forward, it would result in a nose-heavy condition, which could be potentially dangerous on take-off and landing. If the CG is too far aft, the tail-heavy condition will increase the tendency of the aircraft to stall and make landing more difficult.

Stability of the aircraft will also be affected with the CG outside the normal operational limits. Provided the CG lies within specified limits, the aircraft should be safe to fly. The unit of measurement for mass and balance are normally dictated by the aircraft manufacturer and can be either Metric or Imperial terms. Specific definitions for mass and balance ensure they are correctly interpreted.

• Datum: The datum is an imaginary vertical plane from which horizontal measurements are taken. The locations of items such as baggage compartments, fuel tanks, seats and engines are relevant to the datum. There is no fixed rule for the location of the datum. The manufacturer will normally specify the nose of the aircraft, but it could be at the front main bulkhead or even forward of the aircraft nose

• Arm: The horizontal distance from an item or piece of equipment to the datum. The arm's distance is usually measured in inches (or millimetres) and may be preceded by a plus (+) or a minus (-) sign. The plus sign indicates that the distance is aft of the datum and the minus sign indicates distance is forward of the datum

• Moment: The product of a force multiplied by the distance about which the force acts. In the case of mass and balance, the force is the mass (kg/lb) and the distance is the arm (m/in). Therefore, a mass of 40 kilograms, at 3 metres aft of the datum will have a moment of 40 x 3 = 120 kg/m. It is important to consider whether a value is positive (+ve) or negative (-ve) when moments are calculated and the following conventions are used:

Distances horizontal: aft of the datum (+), forward of the datum (-).
Weight: added (+), removed (-).

• Centre of Gravity (CG): This is the point about which all of the mass of the aircraft or object is concentrated. An aircraft could be suspended from this point and it would not adopt a nose-down nor a tail-down attitude.

• Centre of Gravity Balance Limits: For normal operation of the aircraft, the CG should be between the Forward and Aft limits as specified by the manufacturer. If the CG is outside these limits, the aircraft performance will be affected and the aircraft may be unsafe.

• Dry Operating Mass: The total mass of the aeroplane, ready for a specific type of operation, excluding all usable fuel and traffic load. This mass includes crew and crew baggage, catering and removable passenger service equipment, potable water and lavatory chemicals.

• Maximum Zero Fuel Mass: The maximum permissible mass of an aircraft with no usable fuel. Fuel contained in certain tanks must be included if this is explicitly mentioned in the aircraft’s Flight Manual limitations.

• Maximum Structural Take-Off Mass (MTOM): The maximum permissible total aeroplane mass at the start of the take-off run.

• Maximum Structural Landing Mass: The maximum permissible total aeroplane mass upon landing under normal circumstances.

• Traffic Load: This includes the total mass of passengers, baggage and cargo, including any non-revenue load.

Mass and Balance Documentation

The Mass and Balance documentation used by an operator must include certain basic information, which is listed below. Subject to the approval of the authority, some of this information may be omitted.

A. Aeroplane registration and type
B. Flight identification number and date
C. Identity of the commander
D. Identity of the person who prepared the document
E. Dry operating mass and the corresponding CG of the aeroplane
F. Mass of the fuel at take-off and the mass of trip fuel
G. Mass of consumables other than fuel
H. Load components that include passengers, baggage, freight and ballast
I. Take-off Mass, Landing Mass and Zero Fuel mass.
J. The load distribution
K. Aeroplane CG positions
L. Limiting mass and CG values

FREQUENCY OF WEIGHING

Aircraft must be weighed before entering service, to determine the individual mass and CG position. This should be done once all manufacturing processes have been completed. The aircraft must also be re-weighed within four years from the date of manufacture, if individual mass is used, or within nine years from the date of manufacture, if fleet masses are used.

The mass and CG position of an aircraft must be periodically re-established. The maximum interval between one aircraft weigh and the next, must be defined by the operator, but not exceed the four/nine year limits

15.7 CALCULATION OF MASS AND CG OF ANY SYSTEM

The position of the CG of any system (refer to Fig. 1) may be found using the following process:

• Total Mass is calculated, by adding the mass of each load (plus the mass of the beam)

• The moment of each load is calculated, by multiplying the mass by the arm (distance from the reference datum)

• ALL the moments are added together, to provide the Total Moment

• Total Moment is divided by the Total Mass to give CG position.

Aviation Hydraulic Oil

msmprr@yahoo.com (Aviation) — Sat, 04 Sep 2010 12:00:00 +0000

Aircraft hydraulic fluids fall under various specifications:

Common petroleum-based:

Mil-H-5606: Mineral base, flammable, fairly low flashpoint, usable from −65 °F (−54 °C) to 275 °F (135 °C), red color
Mil-H-83282: Synthetic hydrocarbon base, higher flashpoint, self-extinguishing, backward compatible to -5606, red color, rated to −40 °F (−40 °C) degrees.
Mil-H-87257: A development of -83282 fluid to improve its low temperature viscosity.

Phosphate-ester based:

BMS 3-11: Skydrol 500B-4, Skydrol LD-4, Skydrol 5 and Exxon HyJetIV-A plus – Typically light purple, not compatible with petroleum-based fluids, will not support combustion.

Radial Engines

msmprr@yahoo.com (Aviation) — Sat, 04 Sep 2010 11:42:00 +0000

Radial Engines

Nose Section

In general, it is either tapered or round in order to

place the metal under tension or compression instead

of shear stresses.

A tapered nose section is

used quite frequently on direct-drive, low-powered

engines, because extra space is not required to house

the propeller reduction gear.

The nose section on engines which develop from

1,000 to 2,500 hp. is usually rounded and sometimes

ribbed to get as much strength as possible.

Aluminum alloy is the most widely used material

because of its adaptability to forging processes and

its vibration-absorbing characteristics.

Since the nose section transmits many varied

forces to the main or power section, it must be

properly secured to transmit the loads efficiently.

It also must have intimate contact to give rapid and

uniform heat conduction, and be oiltight to prevent

leakage. This is usually accomplished by an offset

or ground joint, secured by studs or capscrews.

On some of the larger engines, a small chamber

is located on the bottom of the nose section to collect

the oil. This is called the nose section oil sump.

Power Section

On engines equipped with a two-piece master rod

and a solid-type crankshaft, the main or power

crankcase section may be solid, usually of aluminum

alloy.

This portion of the engine is often called the

power section, because it is here that the reciprocating

motion of the piston is converted to rotary

motion of the crankshaft.

Because of the tremendous loads and forces from

the crankshaft assembly and the tendency of the

cylinders to pull the crankcase apart, especially in

extreme conditions when a high-powered engine is

detonated, the main crankcase section must be very

well designed and constructed.

The machined surfaces on which the cylinders

are mounted are called cylinder pads. They are

provided with a suitable means of retaining or

fastening the cylinders to the crankcase.

The inner portion of the cylinder pads are sometimes

chamfered or tapered to permit the installation

of a large rubber O-ring around the cylinder

skirt, which effectively seals the joint between the

cylinder and the crankcase pads against oil leakage.

Diffuser Section

The diffuser or supercharger section generally is

cast of aluminum alloy, although, in a few cases,

the lighter magnesium alloy is used.

Because of the elongation and contraction of the

cylinders, the intake pipes which carry the mixture

from the diffuser chamber through the intake valve

ports are arranged to provide a slip joint which

must be leakproof. The atmospheric pressure on

the outside of the case of an unsupercharged engine

will be higher than on the inside, especially when

the engine is operating at idling speed. If the

engine is equipped with a supercharger and operated

at full throttle, the pressure will be considerably

higher on the inside than on the outside of the case.

If the slip joint connection has a slight leakage,

the engine may idle fast due to a slight leaning of

the mixture. If the leak is quite large, it may not

idle at all. At open throttle, a small leak probably

would not be noticeable in operation of the engine,

but the slight leaning of the fuel/air mixture might

cause detonation or damage to the valves and valve

seats.

Accessory Section

The accessory (rear) section usually is of cast

construction, and the material may be either aluminum

alloy, which is used most widely, or magnesium,

which has been used to some extent.

the

increased demands for electric current on large

aircraft and the requirements of higher starting

torque on powerful engines have resulted in an

increase in the size of starters and generators.

This means that a greater number of mounting bolts

must be provided and, in some cases, the entire

rear section strengthened.

In some cases there is a duplication of drives,

such as the tachometer drive, to connect instruments

located at separate stations.

The accessory section provides a mounting place

for the carburetor, or master control, fuel injection

pumps, engine-driven fuel pump, tachometer

generator, synchronizing generator for the engine

analyzer, oil filter, and oil pressure relief valve.

Bleeding a Shock Strut

msmprr@yahoo.com (Aviation) — Fri, 27 Aug 2010 12:41:00 +0000

Bleeding a Shock Strut

If the fluid level of a shock strut has becomeextremely low, or if for any other reason air is
trapped in the strut cylinder, it may be necessary to bleed the strut during the servicing operation. Bleeding is usually performed with the aircraft placed on jacks. In this position the shock struts can be extended and compressed during the filling operation, thus expelling all the entrapped air. The following is a typical bleeding procedure :
Construct a bleed hose containing a fitting suitable for making an airtight connection to the shock strut filler opening. The base should be long enough to reach from the when the aircraft is on jacks.

Jack the aircraft until all shock struts are fully extended.

Release the air pressure in the strut to bebled.

Remove the air valve assembly.

Fill the strut to the level of the filler port with an approved type hydraulic fluid.
Attach the bleed hose to the filler port and insert the free end of the hose into a container of clean hydraulic fluid, making sure that this end of the hose is below the surface of the hydraulic fluid.

Place an exerciser jack or other suitable single-base jack under the shock strut jacking point. Compress and
extend the strut fully by raising and lowering the jack until the flow of air bubbles from the strut has completely stopped.

Compress the strut slowly and allow it to extend by its own weight.

Remove the exerciser jack, and then lower and remove all other jacks.

Remove the bleed hose from the shock strut.
Install the air valve and inflate the strut

msmprr@yahoo.com (Aviation) — Fri, 27 Aug 2010 12:27:00 +0000

Landing Gear System

Undercarriage Configuration
1. Conventiol -Main wheel + tail wheel
2. Tricycle - Main wheel + nose wheel
3. Tandom. - Main wheel + out trigger wheels

Shock Struts
Shock struts are self-contained hydraulic units that support an aircraft on the ground and protect
the aircraft structure by absorbing and dissipating the tremendous shock loads of landing. Shock struts must be inspected and serviced regularly to function efficiently.

SHIMMY DAMPERS
A shimmy damper controls vibration, or shimmy, through hydraulic damping. The damper is either attached to or built integrally with the nose gear and prevents shimmy of the nosewheel during taxiing, landing, or takeoff. There are three types of shimmy dampers commonly used on aircraft:
(1)The piston type,
(2) vane type, and
(3) features incorporated in the nosewheel power steering system of some aircraft.

Steer Damper
A steer damper is hydraulically operaied and accomplishes the two separate functions of steering
and/or eliminating shimmying.

Identification of Aircraft Parts

msmprr@yahoo.com (Aviation) — Thu, 19 Aug 2010 13:04:00 +0000

ICAO phonetic alphabet

msmprr@yahoo.com (Aviation) — Mon, 16 Aug 2010 04:29:00 +0000

ICAO phonetic alphabet :

A ALPHA AL fah

B BRAVO BRAH VO

C CHARLIE CHAR lee

D DELTA DELL tah

E ECHO ECK oh

F FOXTROT FOKS trot

G GOLF GOLF

H HOTEL hoh TELL

I INDIA IN dee ah

J JULIETT JEW lee ETT

K KILO KEY loh

L LIMA LEE mah

M MIKE MIKE

N NOVEMBER no VEM ber

O OSCAR OSS cah

P PAPA pah PAH

Q QUEBEC keh BECK

R ROMEO ROW me oh

S SIERRA see AIR ah

T TANGO TANG go

U UNIFORM YOU nee form

V VICTOR VIK tah

W WHISKEY WISS key

X X-RAY ECKS RAY

Y YANKEE YANG key

Z ZULU ZOO loo