3 Turbine Engines (1)saf

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TURBINE ENGINES

INTRODUCTION

Efforts to design a working gas turbine engine had been under way for years prior to World War II. Engineers even-tually succeeded in placing a few engines in combat aircraft briefly during the closing stages of the war. The wareffort had brought about many advances in gas turbine technology which could now be used for commercial air-craft design. Turbine engines offered many advantages over reciprocating engines and airlines were interested.Increased reliability, longer mean times between overhaul, higher airspeeds, ease of operation at high altitudes,and a high power to engine weight ratio made turbine power very desirable. Aircraft such as Lockheed's SuperConstellation represented the practical limits of piston power technology and required frequent engine mainte-

nance; therefore, air carriers turned to gas turbine engines for solutions. During the decade of the 50's, a gradualtransfer from piston power to gas turbine jets and turboprops started taking place. Old workhorses such as theDouglas DC-3 and DC-7 gave way to the Boeing 707 and Douglas DC-8.


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DESIGN AND CONSTRUCTION

Newton's third law of motion states that for everyaction, there is an equal and opposite reaction. Jet

propulsion applies this law by taking in a quantityof air and accelerating it through an orifice or noz-zle. The acceleration of the air is the action and for-ward movement is the reaction. In nature, a squid

propels itself through the water using a form of jetpropulsion. A squid takes sea water into its bodyand uses its muscles to add energy to the water, thenexpels the water in the form of a jet. This action pro-duces a reaction that propels the squid forward.[Figure 3-1]

Figure 3-1. Many technological developments were madeby observing nature in action. A squid propels itselfthrough the water by jet reaction in much the same way aturbojet engine propels an aircraft.

As early as 250 B.C., a writer and mathematiciannamed Hero devised a toy that used the reaction

principle. The toy, called the aeolipile, consisted ofa covered kettle of water that was heated to producesteam. The steam was then routed through two ver-tical tubes and into a spherical container. Attachedto the spherical container were several discharge

tubes arranged radially around the container. Assteam filled the container, it would escape throughthe discharge tubes and cause the sphere to rotate.[Figure 3-2]

A more modern example of Newton's reaction prin-ciple is observed when the end of an inflated bal-loon is released. As the air in the balloon rushes outthe opening, the balloon flies wildly around a room.In spite of the everyday examples, scientists' effortsto apply Newton's reaction principle to mechanicaldesigns met with little success until this century.

Figure 3-2. Hero's aeolipile, conceived long before theacceptance of Newton's Laws of Motion, proved that power

by reaction was possible.

HISTORY OF JET PROPULSION

The history of mechanical jet propulsion began in1900, when Dr. Sanford Moss submitted his mastersthesis on gas turbines. Later, Dr. Moss became anengineer for the General Electric Company inEngland. While there, Dr. Moss had the opportunityto apply some of his concepts in the development ofthe turbo-supercharger. This unique superchargerconsisted of a small turbine wheel that was driven

by exhaust gases. The turbine was then used to

drive a supercharger.

Research done by Dr. Moss influenced FrankWhittle of England in the development of what

became the first successful turbojet engine. Dr.Whittle was granted his first patent for the jet engine


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Turbine Engines 3-3

Figure 3-3. Dr. Frank Whittle of England patented the first

turbojet engine, the Whittle W1, in 1930. Its first flight

occurred in a Gloster E28/39 aircraft in 1941.

in 1930 and eleven years later, his engine completedits first flight in a Gloster model E28/39 aircraft. Theengine produced about one thousand pounds of

thrust and propelled the aircraft at speeds over 400miles per hour. [Figure 3-3]

While Whittle was developing the gas turbineengine in England, Hans Von Ohain, a German engi-neer, designed and built a jet engine that produced1,100 pounds of thrust. This engine was installed inthe Heinkel He-178 aircraft and made a successfulflight on August 27, 1939. As a result, it becamerecognized as the first practical flight by a jet pro-

pelled aircraft. [Figure 3-4]

In the United States, research in the field of jetpropulsion was lagging. Most of thecountry's

Figure 3-4. German engineer Hans Von Ohain designed and

built the turbojet engine that powered the Heinkel He-178

to the world's first jet-powered flight in 1939.

Figure 3-5. First flown in 1942, the Bell XP-59 was the first

American jet-powered aircraft.

efforts were being directed toward the developmentand production of high powered reciprocatingengines. However, in 1941 the General ElectricCompany received a contract to research anddevelop a gas turbine engine. General Electric waschosen for this important project because of its

extensive experience in building electrical generat-ing turbines and turbo-superchargers. The resultwas the GE-lA engine, a centrifugal-compressortype engine that produced approximately 1,650

pounds of thrust. Two of these engines were used topower the Bell XP-59 "Airacomet" which flew forthe first time in October 1942. The Airacomet

proved the concept of jet powered flight, but wasnever used in combat due to its limited flight timeof 30 minutes. [Figure 3-5]

JET PROPULSION TODAY

Today, the majority of commercial aircraft utilizesome form of jet propulsion. In addition, there arecurrently several manufacturers that produce entirelines of jet powered aircraft that cruise in excess of600 miles per hour and carry more than four hun-dred passengers or several tons of cargo.

Another step in the progression of commercial andmilitary aviation was the ability to produce anengine that would propel an aircraft faster than thespeed of sound. Today, there are several military air-craft that travel at speeds in excess of Mach one.One such aircraft is the SR-71 Blackbird which flysin excess of Mach five. In commercial aviation how-ever, there is currently only one aircraft that fliesfaster than Mach one. This aircraft, the Concorde,was built by the British and French and placed intoservice in the mid seventies. Currently, there aremore than ten Concordes in service that are capableof flying at 2.2 times the speed of sound.

In addition to military and commercial aviation, jetpropulsion has become extremely popular for useon business jets. These two and three engine aircraft

have become extremely popular in recent years duein part to the efficiency and reliability of jet engines.


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3-4 Turbine Engines

TYPES OF JET PROPULSION

Newton's reaction principle has been applied to

several propulsive devices used in aviation. Allproduce thrust in the same manner, they acceleratea mass of gases within the engine. The most com-mon types of propulsive engines are the rocket, theramjet, the pulsejet, and the gas turbine.

ROCKET

A rocket is a nonairbreathing engine that carries itsown fuel as well as the oxygen needed for the fuelto burn. There are two types of rockets in use:solid-propellant rockets and liquid-propellantrockets. Solid-propellant rockets use a solid fuel

that is mixed with an oxidizer and formed into aspecific shape that promotes an optimum burningrate. Once ignited, the fuel produces an extremelyhigh velocity discharge of gas through a nozzle atthe rear of the rocket body. The reaction to therapid discharge is forward motion of the rocket

body. Solid fuel rockets are used primarily topropel some military weapons and, at times,provide additional thrust for takeoff of heavilyloaded aircraft. These booster rockets attach to anaircraft structure and provide the additional thrustneeded for special-condition takeoffs. [Figure 3-6]

The second type of rocket is the liquid-fuel rocket,which uses fuel and an oxidizing agent such as liq-uid oxygen. The two liquids are carried in tanksaboard the rocket. When the liquids are mixed, thereaction is so violent that a tremendous amount ofheat is generated. The resulting high velocity gas jet

behind the rocket provides enough thrust to propelan object.

RAMJET

A ramjet engine is an athodyd, or aero-thermody-

namic-duct. Ramjets are air-breathing engines with

Figure 3-7. As a ramjet moves forward, air enters the intakeand proceeds to a combustion chamber where fuel isadded. Once ignited, the heat from the burning fuel accel-

erates the flow of air through a venturi to produce thrust.

no moving parts. However, since a ramjet has no rotat-ing compressor to draw air into the engine, a ramjetmust be moving forward at a high velocity before itcan produce thrust. Once air enters the engine, fuel isinjected and ignited to provide the heat needed toaccelerate the air and produce thrust. Because ramjetsmust be moving forward to produce thrust, they arelimited in their use. At present, ramjets are used insome military weapons delivery systems where thevehicle is accelerated to a high initial velocity so theramjet can take over for sustained flight. [Figure 3-7]

PULSEJET

Pulsejet engines are similar to ramjets except thatthe air intake duct is equipped with a series of shut-ter valves that are spring loaded to the open

posti-tion. Air drawn through the open valvesenters a combustion chamber where it is heated by

burning fuel. As the air within the combustionchamber expands, the air pressure increases to the

point that the shutter valves are forced closed.Once closed, the expanding air within the chamber

is forced rearward to produce thrust. A pulsejet istypically considered more useful than a ramjetbecause pulsejets will produce thrust prior to beingaccelerated to a high forward speed. [Figure 3-8]

GAS TURBINE ENGINE

The gas turbine engine is by far the most practicalform of jet engine in use today. In fact, the turbineengine has become the standard on nearly all trans-

port category, business, and military aircraft.Because of this, the discussion presented in this sec-tion will focus on the gas turbine engine. The four

most common types of gas turbine engines are theturbojet, turbo-propeller, turboshaft, and turbofan.

Figure 3-6. RATO, or rocket assisted takeoff devices aresmall, solid propellant rocket motors that are attached toan airplane to provide additional thrust for high altitude oroverweight takeoff conditions.


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Turbine Engines 3-5

Figure 3-8. (A) In the pulsejet engine, air is drawn into the

combustion chamber and mixed with fuel when the shutter

valves open. (B) As the fuel burns, the air pressure within

the chamber increases and forces the shutter valves to

close. Once closed, the expanding air within the engine

accelerates rearward through the exhaust nozzle to pro-

duce thrust.

TURBOJET ENGINES

The basic operating principles of a turbojet engine

are relatively straight forward; air enters through aninlet duct and proceeds to the compressor where itis compressed. Once compressed, the air flows tothe combuster section where fuel is added andignited. The heat generated by the burning fuelcauses the compressed air to expand and flowtoward the rear of the engine. As the air moves rear-ward, it passes through a set of turbine wheels thatare attached to the same shaft as the compressor

blades. The expanding air spins the turbines which,in turn, drives the compressor. Once past the tur-

bines, the air proceeds to exit the engine at a muchhigher velocity than the incoming air. It is this dif-

ference in velocity between the entering and exitingair that produces thrust.

When discussing a turbojet engine you must befamiliar with the term engine pressure ratio, orEPR. An engine's EPR is the ratio of the turbine dis-charge pressure to the engine inlet air pressure. EPRgauge readings are an indication of the amount ofthrust being produced for a given power lever set-ting. Total pressure pickups, or EPR probes, mea-sure the air pressure at two points in the engine; oneEPR probe is located at the compressor inlet and asecond EPR probe is located just aft of the last stage

turbine in the exhaust section. EPR readings areoften used as verification of power settings for take-

off, climb, and cruise. EPR readings are affected byand are dependent on pressure altitude and outsideair temperature (OAT).

TURBOPROP ENGINES

A gas turbine engine that delivers power to a pro-peller is referred to as a turboprop engine.Turboprop engines are similar in design to turbojetengines except that the power produced by a turbo-

prop engine is delivered to a reduction gear systemthat spins a propeller. Reduction gearing is neces-sary in turboprop engines because optimum pro-

peller performance is achieved at much slowerspeeds than the engine's operating rpm. Turbopropengines are used extensively in business and com-

muter type aircraft because the combination of jetpower and propeller efficiency provides good per-formance characteristics at speeds between 300 and400 miles per hour. In addition, most turbopropengines provide the best specific fuel consumptionof any gas turbine engine. [Figure 3-9]

TURBOSHAFT ENGINES

A gas turbine engine that delivers power to a shaftthat can drive something else is referred to as a tur-

boshaft engine. The biggest difference between aturbojet and turboshaft engine is that on a turboshaft

engine, most of the energy produced by the expand-ing gases is used to drive a turbine rather than pro-duce thrust. Many helicopters use a turboshaft typeof gas turbine engine. In addition, turboshaftengines are widely used as auxiliary power unitsand in industrial applications to drive electricalgenerators and surface transportation systems.Output of a turboprop or turboshaft engine is mea-sured by shaft horsepower rather than thrust.

TURBOFAN ENGINES

A turbofan engine consists of a multi-bladed ducted

propeller driven by a gas turbine engine. Turbofanswere developed to provide a compromise between

Figure 3-9. Turboprop powerplants have become a popular

choice on corporate twin-engine aircraft.


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3-6 Turbine Engines

Figure 3-10. (A) A forward-fan turbofan engine uses a relatively large diameter ducted fan that produces thrust and providesintake air to the compressor. (B) An aft-fan turbofan engine has a fan mounted on the aft turbine. This arrangement is rarelyused, since an aft fan cannot contribute to air compression at the inlet.

the best features of the turbojet and the turboprop.Turbofan engines have turbojet-type cruise speedcapability, yet retain some of the short-field takeoffcapability of a turboprop. Nearly all present day air-liners are powered by turbofan engines for the rea-sons just mentioned as well as the fact thatturbo-fans are very fuel efficient.

A turbofan engine may have the fan mounted toeither the front or back of the engine. Engines thathave the fan mounted in front of the compressor arecalled forward-fan engines, while turbofan enginesthat have the fan mounted to the turbine section arecalled aft-fan engines. [Figure 3-10]

The inlet air that passes through a turbofan engineis usually divided into two separate streams of air.

One stream passes through the engine core while asecond stream coaxially bypasses the engine core. Itis this bypass stream of air that is responsible for the

term bypass engine. When discussing bypassengines there are three terms you must be familiarwith; they are thrust ratio, bypass ratio, and fan

pressure ratio. A turbofan engine's thrust ratio is acomparison of the thrust produced by the fan to thethrust produced by the engine core exhaust. On theother hand, a turbofan's bypass ratio refers to theratio of incoming air that bypasses the core to theamount of air that passes through the engine core.Turbofans in civil aircraft are generally divided intothree classifications based on bypass ratio:

1.Low bypass (1:1)2.Medium bypass (2:1 or 3:1)3.High bypass (4:1 or greater)


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Turbine Engines 3-7

Figure 3-11. (A) Bypass air is ejected directly overboardin forward-fan engines with a short fan duct. (B)However, in a ducted fan, bypass air is ducted along theengine's entire length.

Generally, airflow mass in the fan section of a lowbypass engine is the same as airflow mass in thecompressor. The fan discharge could be slightlyhigher or lower depending on the engine model, but

bypass ratios are approximately 1:1. In someengines the bypass air is ducted directly overboardthrough a short fan duct. However, in a ducted fanengine, the bypass air is ducted along the entirelength of the engine. Full fan ducts reduce aerody-namic drag and noise emissions. In either case, theend of the duct usually has a converging dischargenozzle that increases velocity and produces reactivethrust. [Figure 3-11]

Medium or intermediate bypass engines have air-flow bypass ratios ranging from 2:1 to 3:1. Theseengines have thrust ratios similar to their bypass

ratios. The fans used on these engines have a largerdiameter than the fans used on low bypass enginesof comparable power. Fan diameter determines afan's bypass ratio and thrust ratio.

High bypass turbofan engines have bypass ratios of4:1 or greater and use the largest diameter fan of anyof the bypass engines. High bypass turbines offerhigher propulsive efficiencies and better fuel econ-omy than low or medium bypass turbines.Consequently, they are the engines of choice onlarge airliners used for long flights. Some commonhigh bypass turbofan engines include Pratt and

Whitney's JT9D and PW4000, the Rolls-RoyceRB-211, and the General Electric CF6. Oneversion of

Figure 3-12. All high bypass turbofan engines use largediameter fans that produce bypass ratios of 4:1 or greater.

the JT9D has a bypass ratio of 5:1 with 80 percent ofthe thrust provided by the fan, and only 20 percent

by the core engine. [Figure 3-12]

Another term you must be familiar with is fan pres-sure ratio which is the ratio of air pressure leavingthe fan to the air pressure entering the fan. The fan

pressure ratio on a typical low bypass fan is approx-imately 1.5:1, whereas for some high bypass fans thefan pressure ratio may be as high as 7:1. To obtainhigh fan pressure ratios, most high bypass enginesare designed with high aspect ratio blades. Aspectratio is the ratio of a blade's length to its width, orchord. Therefore, a long blade with a narrow chordhas a higher aspect ratio than a short blade with awide chord. Although high aspect ratio fan bladesare used most often, low aspect ratio blades arecoming into wider use today. Technologicaladvances in blade construction have overcome theweight problems associated with low aspect ratio

blades in the past. Weight savings in low aspectratio blades have been achieved with hollow tita-nium blades having composite inner reinforcementmaterials. Additionally, low aspect ratio blades are

desirable because of their resistance to foreignobject damage, especially bird strikes.

UNDUCTED FAN ENGINES

Recent developments have produced new enginedesigns with higher efficiencies than anything cur-rently in use. The new engines are designated ultrahigh bypass (UHB) propfan and unducted fanengine (UDF). These new designs utilize titanium,lightweight stainless steel, and composite materialsto surpass the fuel economy of several high bypassturbofan engines by more than 15 percent. Engine

designers have achieved 30:1 bypass ratios by incor-porating single or dual propellers with composite


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3-70 Turbine Engines

Figure 3-18. The Hawker-Siddeley ou i "Nimrod" was devel-

oped from the de Havilland Comet airframe and utilizes

wing mounted air inlets that are aerodynamicaliy shaped to

reduce drag.

WING-MOUNTED INLETS

Some aircraft with engines mounted inside the wingsfeature air inlet ducts in the wing's leading edge.Aircraft such as the Aerospatiale Caravelle, deHavilland Comet, and de Havilland Vampire all utilizewing-mounted inlets. Typically, wing-mounted inletsare positioned near the wing root area. [Figure 3-18]

FUSELAGE-MOUNTED INLETS

Engines mounted inside a fuselage typically use airinlet ducts located near the front of the fuselage. For

example, many early military aircraft were designedwith an air inlet duct in the nose of the fuselage. Inaddition, some modern supersonic military aircrafthave inlet ducts located just under the aircraft nose.Although using an air inlet of this type allows theaircraft manufacturer to build a more aerodynamicaircraft, the increased length of the inlet does intro-duce some inefficiencies. [Figure 3-19]

Some military aircraft use air inlet ducts mountedon the sides of the fuselage. This arrangement workswell for both single and twin engine aircraft. Bymounting an intake on each side of an aircraft, the

duct length can be shortened without adding asignificant amount of drag to the aircraft. However,a

Figure 3-19. The single-enframe miei: uuu takes full advan-

tage of ram effect much like engine-mounted air inlets.Although the aircraft is aerodynamicaliy clean, the length of

the duct makes it slightly less efficient than

engine-mounted types.

Figure 3-20. The divided-entrance duct with side-mounted

intakes has a shorter length, providing improved inlet

efficiency.

disadvantage to this arrangement is that somesudden flight maneuvers can cause an imbalance in

ram air pressure between the two intakes. The airpressure imbalance felt on the compressor faceresults in a slight loss of power. [Figure 3-20]

SUBSONIC INLETS

A typical subsonic air inlet consists of a fixed geom-etry duct whose diameter progressively increasesfrom front to back. This divergent shape works likea venturi in that as the intake air spreads out, thevelocity of the air decreases and the pressureincreases. This added pressure contributes signifi-

cantly to engine efficiency once the aircraft reachesits design cruising speed. At this speed, the com-pressor reaches its optimum aerodynamic efficiencyand produces the most compression for the best fueleconomy. It is at this design cruise speed that theinlet, compressor, combustor, turbine, and exhaustduct are designed to match each other as a unit. Ifany section mismatches any other because of dam-age, contamination, or ambient conditions, engine

performance suffers. For additional information onsubsonic air inlets, refer to the discussion on turbineengine induction systems in Section B of Chapter 5.

SUPERSONIC INLETS

On supersonic aircraft a typical air inlet duct haseither a fixed or variable geometry whose diameter

progressively decreases, then increases from front toback. This convergent-divergent shape is used toslow the incoming airflow to subsonic speed beforeit reaches the compressor.

In addition to the convergent-divergent shape,many supersonic inlets employ a movable plug orthroat that changes the duct's geometry. The vari-able geometry is necessary so the duct can be

adjusted as needed to accomodate a wide range offlight speeds. For additional information on super-


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Figure 3-23. When a pilot actuates this type of sand separator, a small vane extends into the airstream. The inertia of the sand and

ice particles after they pass through the venturi carries them past the air intake and discharges them overboard.

Another type of separator used on some turbopropaircraft incorporates a movable vane which extendsinto the inlet airstream. Once extended, the vanecreates a more prominent venturi and a sudden turnin the engine inlet. Combustion air can follow thesharp curve but sand or ice particles cannot becauseof their inertia. The movable vane is operated by a

pilot through a control handle in the cockpit.[Figure 3-23]

Some gas turbine engine inlets have a tendency toform a vortex between the ground and the inlet dur-ing ground operations. This vortex can become

strong enough to lift water and debris such as sand,small stones, or small hardware from the groundand direct it into the engine. To help alleviate this

problem, a vortex dissipater, sometimes called avortex destroyer or blow-away jet is installed onsome gas turbine engines. A typical vortex dissipa-ter routes high pressure bleed air to a discharge noz-zle located in the lower part of the engine cowl.This discharge nozzle directs a continuous blast of

bleed air between the ground and air inlet to pre-vent a vortex from developing. Most aircraftequipped with a vortex dissipater also have a land-ing gear switch that arms the dissipater wheneverthe engine is operating and weight is on the maingear. [Figure 3-24]

COMPRESSOR SECTIONAs discussed earlier, a gas turbine engine takes in aquantity of air, adds energy to it, then discharges theair to produce thrust. Based on this, the more airthat is forced into an engine, the more thrust theengine can produce. The component that forces airinto the engine is the compressor. To be effective, amodern compressor must increase the intake air

pressure 20 to 30 times above the ambient air pres-sure and move the air at a velocity of 400 to 500 feet

Figure 3-24. Engines that utilize a vortex dissipater use high

pressure bleed air from the compressor to prevent the for-mation of a low pressure vortex that can suck debris into

the engine.


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Turbine Engines 3-13

per second. One way of measuring a compressor'seffectiveness is to compare the static pressure of thecompressor discharge with the static air pressure at

the inlet. If the discharge air pressure is 30 timesgreater than the inlet air pressure, that compressorhas a compressor pressure ratio of 30:1.

In addition to supporting combustion and providingthe air necessary to produce thrust, the compressorsection has several secondary functions. For exam-

ple, a compressor supplies bleed air to cool the hotsection and heated air for anti-icing. In addition,compressor bleed air is used for cabin

pressurization, air conditioning, fuel systemdeicing, and pneumatic engine starting.There aretwo basic types of compressors used today; the

centrifugal flow compressor and the axial flowcompressor. Each is named according to thedirection the air flows through the compressor, andone or both may be used in the same engine.

CENTRIFUGAL FLOW COMPRESSORS

The centrifugal compressor, sometimes called aradial outflow compressor, is one of the earliestcompressor designs and is still used today in somesmaller engines and auxiliary power units (APU's).Centrifugal compressors consist of an impeller, a

diffuser, and a manifold.[Figure 3-25]

The impeller, or rotor, consists of a forged disk withintegral blades, fastened by a splined coupling to acommon power shaft. The impeller's function is totake air in and accelerate it outward by centrifugalforce.Centrifugal compressors can have one or twoimpellers. Compressors having only one impellerare referred to as single-stage compressors whilecompressors having two impellers are referred to asdouble-stage compressors. Although a two-stage

Figure 3-26. A two-stage impeller is sometimes used toobtain higher compressor pressure ratios. However, due toefficiency losses, centrifugal compressors typically do notexceed two stages.

impeller compresses the air more than a single-stageimpeller, the use of more than two stages in a com-

pressor is typically considered impractical. Thebenefits of additional stages are negated by theenergy lost when the airflow slows down as it

passes from one impeller to the next. In addition,the added weight from each additional impellerrequires more energy from the engine to drive the

compressor. [Figure 3-26]

When two impellers are mounted back-to-back adouble-sided or double-entry impeller is created. Asingle-stage, double-sided impeller allows a highermass airflow than that of a similar sizedsingle-stage, single-sided impeller. Therefore,engines "with double-sided impellers typicallyhave a smaller overall diameter. [Figure 3-27]

Figure 3-25. A single-stage centrifugal compressor consistsof an impeller, a diffuser, and a compressor manifold.

Figure 3-27. A single-stage, dual-sided impeller enables asmall diameter engine to produce a high mass airflow.


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One drawback of the double-sided impeller is thatthe ducting required to get the intake air from oneside of the impeller to the other is complicated. For

example, included in the ducting for double-entrycompressor engines is a plenum chamber. Thischamber is necessary because the air must enter theengine at almost right angles to the engine axis.Therefore, in order to give a positive flow, the airmust surround the engine compressor at a positive

pressure before entering the compressor. In additionto the plenum chamber, some double-entry com-

pressors utilize auxiliary air-intake doors (blow-indoors). These blow-in doors admit air into theengine compartment during ground operation whenair requirements for the engine exceed that of theincoming airflow. The doors are held closed by

spring action when the engine is not operating.During operation, however, the doors open when-ever engine compartment pressure drops belowatmospheric pressure. During takeoff and flight, ramair pressure in the engine compartment aids thesprings in holding the doors closed.

Once through the impeller, the air is expelled into adivergent duct called a diffuser, where it losesvelocity and increases in pressure. The diffuser actsas a divergent duct where the air spreads out, slowsdown, and increases in static pressure.

The compressor manifold distributes the air in asmooth flow to the combustion section. The mani-fold has one outlet port for each combustion cham-

ber so that the air is evenly divided. A compressoroutlet elbow is bolted to each of the outlet ports.The elbows act as air ducts and are often referred toas outlet ducts, outlet elbows, or combustion cham-ber inlet ducts. These outlet ducts change the radialdirection of the airflow to an axial direction. To helpthe elbows perform this function in an efficientmanner, turning vanes or cascade vanesare some-

times fitted inside the elbows. These vanes reduceair pressure losses by presenting a smooth, turningsurface. [Figure 3-28]

Centrifugal flow compressors offer several advan-tages including simplicity of manufacture, rela-tively low cost, low weight, low starting powerrequirements, and operating efficiency over a widerange of rotational speeds. In addition, a centrifugalflow compressor's short length and spoke-likedesign allow it to accelerate air rapidly and imme-diately deliver it to the diffuser in a short distance.

Tip speeds of centrifugal compressors may reachMach 1.3, but the pressure within the compressor

Figure 3-28. The turning vanes in a compressor mani-

fold help direct the compressor outlet air to the com-bustion section.(90)

casing prevents airflow separation and provides ahigh transfer of energy into the airflow. Althoughmost centrifugal compressors are limited to twostages, the high pressure rise per stage allows mod-ern centrifugal compressors to obtain compressor

pressure ratios of 15:1.

A typical centrifugal compressor has a few disad-vantages that make it unsuitable for use in some

engines. For example, the large frontal area requiredfor a given airflow increases aerodynamic drag.Also, practical limits on the number of stagesrestrict its usefulness when designing larger andmore powerful engines.

AXIAL FLOW COMPRESSORS

An axial flow compressor has two main elements, arotor and a stator. The rotor consists of rows of

blades fixed on a rotating spindle. The angle andairfoil contour of the blades forces air rearward in

the same manner as a propeller.The stator vanes, onthe other hand, are arranged in fixed rows betweenthe rows of rotor blades andact as diffusersat eachstage, decreasing air velocity and raising pressure.Each consecutive row of rotor blades and statorvanes constitutes a pressure stage. The number ofstages is determined by the amount of air and total

pressure rise required.( )

Unlike a centrifugal compressor, which is capable ofcompressor pressure ratios of 15:1, a single stage inan axial flow compressor is capable of producing acompressor pressure ratio of only 1.25:1. Therefore,high compressor pressure ratios are obtained byadding more compressor stages.


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Figure 3-29. In an axial compressor, airflow velocity is main-

tained nearly constant while air pressure increases as the

airflow proceeds through each stage of compression.

The task of an axial compressor is to raise air pres-sure rather than air velocity. Therefore, each com-

pressor stage raises the pressure of the incoming airwhile the air's velocity is alternately increased thendecreased as airflow proceeds through the compres-sor. The rotor blades slightly accelerate the airflow,then the stator vanes diffuse the air, slowing it andincreasing the pressure. The overall result isincreased air pressure and relatively constant airvelocity from compressor inlet to outlet. [Figure 3-29]

As air passes from the front of an axial flow compressor to the rear, the space between the rotor shaftand the stator casing gradually decreases. Thisshape is necessary to maintain a constant air velocity as air density increases with each stage of com

pression. To accomplish the convergent shape, eachstage of blades and vanes is smaller than the one

preceding it.[Figure3-30]

The case on most axial flow compressors is hori-zontally divided into two halves, allowing the

Figure 3-30. In an axial flow compressor, the divergentshape allows the air velocity to remain nearly constant,

while pressure gradually increases.

removal of one of the halves for inspection or main-tenance of both rotor blades and stator vanes. Thecompressor case also provides a means of extracting

bleed air for ancillary functions.

Some disadvantages of axial flow compressors arerelatively high weight and high starting powerrequirements. Also, the low pressure rise per stage of1.25:1 requires many stages to achieve high compres-sor pressure ratios. Furthermore, axial flow compres-sors are expensive and difficult to manufacture.()

In spite of the disadvantages just mentioned, axialflow compressors outperform centrifugal flow com-

pressors in several areas. High ram efficiency is

obtained because of their straight-through design,which takes full advantage of any ram effect.Another advantage of axial flow compressors istheir ability to obtain higher compressor pressureratios by adding additional stages.In addition, thesmall frontal area of an axial flow compressor helpsto reduce aerodynamic drag.

Compressor Rotor Blades

The rotor blades used in an axial flow compressorhave an airfoil cross-section with a varying angle ofincidence, or twist. This twist compensates for the

blade velocity variation caused by its radius. Inother words, the further from the axis of rotation a

blade section is, the faster it travels. [Figure 3-31]

Figure 3-31. Compressor rotor blades are twisted to compen-

sate for blade velocity variations along the length of the blade.


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Figure 3-32. A dovetail on the base of this compressor blade

fits loosely into a dovetail slot in the compressor wheel. A

small locking device such as a pin, key, or plate prevents

the blade from backing out.

Axial flow compressors typically have 10 to 18 com-pression stages, and in the turbofan engine, the fanis considered to be the first stage rotor. The base, orroot of a rotor blade often fits loosely into the rotordisk. This loose fit allows for easy assembly andvibration damping. As the compressor rotor rotates,

centrifugal force keeps the blades in their correctposition, and the airstream over each blade providesa shock absorbing or cushioning effect. Rotor bladeroots are designed with a number of different shapessuch as a bulb, fir tree, or dovetail. To prevent a

blade from backing out of its slot, most methods ofblade attachment use a pin and a lock tab or lockerto secure the coupling. [Figure 3-32]

Some long fan blades have a mid-span shroud thathelps support the blades,making them more resis-tant to the bending forces created by the airstream.

The shrouds, however, do block some of the airflowand create additional aerodynamic drag thatreduces fan efficiency. In addition, when the mat-ing surfaces on a mid-span shroud become exces-sively worn, the shrouds can overlap. This isknown as shingling and can cause fan vibration andengine damage.

Some blades are cut off square at the tip and arereferred to as flat machine tips. Other blades have areduced thickness at the tips and are called profiletips. All rotating machinery has a tendency tovibrate, and profiling a compressor blade increases

its natural vibration frequency. By increasing theblade's natural frequency above the frequency ofrotation, a blades's vibration tendency is reduced. In

addition, the thin trailing edge of profile tippedblades causes a vortex which increases air velocityand helps prevent air from spilling back over the

blade tips.

On some newer engines the profile tipped bladesare designed with tight running clearances androtate within a shroud strip of abradable material.Since rotor blades are usually made of a stainlesssteel alloy, the shroud strip wears away with no lossof blade length if contact loading takes place.Sometimes after engine shutdown, a high pitchednoise can be heard as the rotor coasts to a stop. Thenoise is caused by contact between the blade tip andshroud strip and is the reason why profile tip bladesare sometimes referred to as squealer tips.

Another blade design that increases compressorefficiency utilizes a localized increase in blade cam-

ber, both at the blade tip and blade root. The pur-pose of this design is to compensate for the frictioncaused by the boundary layer of air near the com-

pressor case. The increased blade camber helpsovercome the friction and makes the blade extremi-ties appear as if they were bent over at each corner,hence the term end bend. [Figure 3-33]

Figure 3-33. End bend refers to the increased blade camberon some compressor blades. The increased camber helps

prevent airflow stagnation near the blade tips.


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Compressor Stator Vanes

Stator vanes are the stationary blades locatedbetween each row of rotating blades in an axial flow

compressor. As discussed earlier, the stator vanesact as diffusers for the air coming off the rotor,decreasing its velocity and raising its pressure. Inaddition, the stators help prevent swirling anddirect the flow of air coming off each stage to thenext stage at the appropriate angle. Like rotor

blades, stator vanes have an airfoil shape. In addi-tion, the angle of attack of stator vanes can be fixedor variable.Stator vanes are normally constructedout of steel or nickel because those metals have highfatigue strength. However, titanium may also beused for stator vanes in the low pressure and tem-

perature stages.

Stator vanes may be secured directly to the com-pressor casing or to a stator vane retaining ring,which is secured to the compressor case. Most sta-tor vanes are attached in rows with a dovetailarrangement and project radially toward the rotoraxis. Stator vanes are often shrouded at their tips tominimize vibration tendencies. [Figure 3-34]

The set of stator vanes immediately in front of the firststage rotor blades are called inlet guide vanes.Thesevanes direct the airflow into the first stage rotor

blades at the best angle while imparting a swirlingmotion in the direction of engine rotation.This actionimproves the aerodynamics of the compressor byreducing the drag on the first stage rotor blades.Someaxial compressors with high compressor pressureratios utilize variable inlet guide vanes plus severalstages of variable stator vanes. These variable inletguide vanes and stators automatically repositionthemselves to maintain proper airflow through theengine under varying operating conditions.

The last set of vanes the compressor air passes throughis the outlet vane assembly. These vanes straighten

Figure 3-34. Compressor stator vanes may be attacheddirectly to the compressor case (A) or to a retaining ringthat is attached to the case by a retaining screw (B). Inaddition, stator vanes are sometimes equipped withshrouds to minimize the effects of vibration.IGVdirectflow to the 1

ststage rotor at best angle

the airflow and eliminate any swirling motion or tur-bulence. The straightened airflow then proceeds to thediffuser to prepare the air mass for combustion.

MULTIPLE-SPOOL COMPRESSORS

In a basic axial flow compressor, the compressor andturbine are connected by a single shaft and rotate as asingle unit. Since there is only one compressor unit,the compressor is commonly referred to as asingle-spool compressor. While single-spoolcompressors are relatively simple and inexpensive tomanufacture, they do have a few drawbacks. Forexample, in a long axial compressor the rear stagesoperate at a fraction of their capacity, while theforward stages are typically overloaded. Furthermore,the large mass of a single-spool compressor does not

respond quickly to abrupt control input changes.[Figure 3-35]

Figure 3-35. In a single-spool compressor, there is only one compressor unit that is connected by a shaft to the turbine section.Rear stages operate at a fraction of capacity while FWD stage overload..Not respond quickly to abrupt control input change


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Figure 3-36. In a dual-spool axial flow engine, the low pressure compressor is driven by the low pressure turbine while the highpressure turbine drives the high pressure compressor. Splitting the compressor creates two rotating groups, each with consider-ably less mass than a single-spool compressor. The smaller mass allows the compressors to respond more quickly to power leverinputs and perform better at high altitudes. In addition, a smaller starter can be used since it turns less mass. (small mass allowcompressor to respond quickly and easy to turn for starting)N2 speed is held relatively constant by FCU GovernorN2 speed up or slow down by ALTand FLT maneuvering

Engine designers devised a way to overcome thelimitations of single-spool compressors by splitting

the compressor into two or three sections. Each section is connected to a portion of the turbine section

by shafts that run coaxially, one within the other.For example, split-compressor engines with twocompressor sections are identified as dual-spool ortwin-spool compressors. The front section of a dual-spool compressor is called the low pressure, lowspeed, or N1compressor. This low pressure com

pressor is typically driven by a two-stage turbine atthe rear of the turbine section. The second compressor section of a twin-spool compressor is called thehigh pressure, high speed, or N2compressor and is

typically driven by a single stage high-pressure turbine at the front of the turbine section. The shaftconnecting the low pressure compressor and tur

bine typically rotates inside the shaft connecting thehigh pressure compressor and turbine. On some tur-

bofan engines, the forward fan is attached to the lowpressure compressor, and they both turn at the samespeed. [Figure 3-36]

Since the spools are not physically connected to oneanother, each is free to seek its own best operatingspeed. However, for any given power lever setting,the high pressure compressor speed is held rela-

tively constant by the fuel control governor. With aconstant energy level at the turbine, the low pres-sure compressor speeds up or slows down withchanges in the inlet air flow caused by atmospheric

pressure fluctuations or flight maneuvering. Forexample, low pressure compressors speed up as the

aircraft gains altitude, since the atmosphere is lessdense and more rotational speed is needed to forcethe required amount of air through the engine.Conversely, as the aircraft descends, the air becomesmore dense and easier to compress so the low pres-sure compressor slows down. This way, the low

pressure compressor supplies the high pressurecompressor with a fairly constant air pressure andmass airflow for each power setting.

On many turbofan engines, the compressor sectionis divided into three sections and is referred to as a

triple-spool compressor. In this arrangement the fanis referred to as the low speed, or N1compressor.The compressor next in line is called the intermedi-ate, or N2compressor, and the innermost compres-sor is the high pressure, or N3compressor. The lowspeed compressor is typically driven by a multiplestage low pressure turbine, while the intermediateand high pressure compressors are driven by singlestage turbines. [Figure 3-37]

COMPRESSOR STALL

As discussed earlier, compressor blades are actually

small airfoils and therefore, are subject to the sameaerodynamic principles that apply to aircraft wings.Like a wing, a compressor blade has an angle ofattack, which is the acute angle between the chord


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Figure 3-37. The triple-spool compressors used on many turbofan engines allow each compressor section to reach its optimumspeed for varying power requirements and flight conditions.(compressor stall turbulence air in inlet abrupt FLTmaneuver abnormal acceltin or decell direction or damage on comp .turbine blade or vane)pulsating or fluttering sound Fluction RPM EGT raise

Air velocity and compressor rotational velocity combine vector which define the AOA(angel of attack)

of the blade and the relative wind. The angle ofattack of a compressor blade is the result of inlet airvelocity and the compressor's rotational velocity.These two forces combine to form a vector, whichdefines the airfoil's actual angle of attack to theapproaching inlet air. As with an aircraft wing, acompressor blade's angle of attack can be changed.

A compressor stall can be described as an imbal-ance between the two vector quantities, inlet veloc-ity, and compressor rotational speed. Compressorstalls occur when the compressor blades' angle of

attack exceeds the critical angle of attack. At thispoint, smooth airflow is interrupted and turbulenceis created with pressure fluctuations. Compressorstalls cause air flowing in the compressor to slowdown and stagnate, sometimes reversing direction.A compressor stall can usually be heard as a pulsat-ing or fluttering sound in its mildest form to a loudexplosion in its most developed state. Quite oftenthe cockpit gauges will not show a mild or transientstall but will indicate a developed stall. Typicalinstrument indications include fluctuations in rpmand an increase in exhaust gas temperature. Mosttransient stalls are not harmful to the engine and

often correct themselves after one or two pulsations.However, severe stalls, or hung stalls, can signifi-cantly impair engine performance, cause loss of

power, and can damage the engine.

The only way to overcome a stalled condition is toreduce the angle of attack on the rotor blades. One

way this can be done is through the use of variableinlet guide vanes( ) and variablestator vanes which direct the incoming air into therotor blades at an appropriate angle. For example,as a compressor's rotational speed decreases, thestator vanes are progressively closed to maintainthe appropriate airflow angle to the proceedingrotor blades. The position of the stator vanes iscontrolled automatically by the fuel control unit.To do this, the fuel control unit monitorscompressor inlet temperature and engine speed.

Another way the angle of attack can be changed isby bleeding off some of the air pressure within thecompressor. To do this, some engines incorporateautomatic air-bleed valves which operate duringlow rpm conditions or during engine startup. Theautomatic valves open to relieve pressure caused byair piling up at the compressor's high pressure end.This regulation of air pressure helps prevent thecompressor from stalling and allows for easierengine starting.

Compressor stalls typically occur when the engineinlet air becomes turbulent or disrupted when an

aircraft flys in severe turbulence or performs abruptflight maneuvers. Another cause is excessive fuelflow produced by a sudden engine acceleration,accompanied by incompatible engine rpm and air-flow combinations. In addition, contamination ordamage to compressor blades, stator vanes, or tur-

bine components can also cause a compressor stall.


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COMBINATION COMPRESSORS

Hybrid axial flow-centrifugal flow compressorswere developed to combine the best features of cen-trifugal and axial compressors and eliminate someof their disadvantages. This design is currently

being used in some smaller engines installed onbusiness jets and helicopters. [Figure 3-38]

COMPRESSOR AIR BLEEDS

In addition to supplying air for combustion, thecompressor supplies high pressure, high tempera-ture air for various secondary functions such ascabin pressurization, heating, and cooling. Also,compressor air is used for deicing, anti-icing, and

for pneumatic engine starting. This air is referred toas bleed air, or customer bleed air and is tappedfrom the compressor through bleed ports at variousstages. A bleed port is a small opening adjacent tothe compressor stage selected for bleed air supply.The choice of which compressor stage to bleed airfrom depends on the air pressure or temperaturerequired for a particular function. Air bled from the

final or highest pressure stage often requires cool-ing, since compression can heat the air to tempera-tures in excess of 650 degrees Fahrenheit.

Bleeding air from the compressor does cause a smallbut noticeable drop in engine power. Sometimespower loss can be detected by observing the enginepressure ratio (EPR) indicator. For example, select-ing the engine inlet anti-ice function causes a dropin EPR and engine rpm if the engine power lever isleft in a fixed position. Exhaust gas temperature(EGT) readings may shift noticeably as well.

DIFFUSER

As air leaves an axial flow compressor and moves

toward the combustion section, it is traveling atspeeds up to 500 feet per second. This is far too fastto support combustion, therefore the air velocitymust be slowed significantly before it enters thecombustion section. The divergent shape of adif-fuser slows compressor discharge while, at thesame time, increasing air pressure to its highestvalue in the engine. The diffuser is usually aseparate section

Figure 3-38. The Garrett TFE731 engine has a two-stage compressor that uses an axial flow compressor for the low pressure stageand a single stage centrifugal compressor for the high pressure stage.


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Figure 3-39. High-velocity air from the compressor section

enters the diffuser where air velocity decreases and air

pressure increases to a maximum.

bolted to the rear of the compressor case and aheadof the combustion section. [Figure 3-39]

COMBUSTION SECTION

A combustion section is typically located directlybetween the compressor diffuser and turbine sec-tion. All combustion sections contain the same

basic elements: one or more combustion chambers(combustors), a fuel injection system, an ignitionsource, and a fuel drainage system.

The combustion chamber or combustor in a turbineengine is where the fuel and air are mixed and

burned. A typical combustor consists of an outer cas-ing with a perforated inner liner. The perforationsare various sizes and shapes, all having a specificeffect on the flame propagation within the liner.

The fuel injection system meters the appropriateamount of fuel through the fuel nozzles into the com-

bustors. Fuel nozzles are located in the combustionchamber case or in the compressor outlet elbows.Fuel is delivered through the nozzles into the liners ina finely atomized spray to ensure thorough mixingwith the incoming air. The finer the spray, the morerapid and efficient the combustion process.

A typical ignition source for gas turbine engines isthe high-energy capacitor discharge system, con-sisting of an exciter unit, two high-tension cables,

and two spark igniters. This ignition system pro-duces 60 to 100 sparks per minute, resulting in a

ball of fire at the igniter electrodes. Some of thesesystems produce enough energy to shoot sparks sev-eral inches, so care must be taken to avoid a lethalshock during maintenance tests.

A fuel drainage system accomplishes the importanttask of draining the unburned fuel after engineshutdown. Draining accumulated fuel reduces the

possibility of exceeding tailpipe or turbine inlettemperature limits due to an engine fire after shut-down. In addition, draining the unburned fuelhelps to prevent gum deposits in the fuel manifold,nozzles, and combustion chambers which arecaused by fuel residue.

To accomplish the task of efficiently burning thefuel/air mixture a combustion chamber must

1.mix fuel and air effectively in the best ratio forgood combustion.

2.burn the mixture as efficiently as possible.3.cool the hot combustion gases to a temperature

the turbine blades can tolerate.

4.distribute hot gases evenly to the turbinesection.

In order to allow the combustion section to mix theincoming fuel and air, ignite the mixture, and coolthe combustion gases, airflow through a combustoris divided into primary and secondary paths.Approximately 25 to 35 percent of the incoming airis designated as primary while 65 to 75 percent

becomes secondary. Primary, or combustion air, isdirected inside the liner in the front end of a com-

bustor. As this air enters the combustor, it passesthrough a set of swirl vanes, which gives the air a

radial motion and slows down its axial velocity toabout five or six feet per second. The reduction inairflow velocity is very important becausekerosene-type fuels have a slow flame propagationrate. Therefore, an excessively high velocityairflow could literally blow the flame out of theengine. This malfunction is known as a flameout. Avortex created in the flame area provides theturbulence required to properly mix the fuel andair. Once mixed, the combustion process iscomplete in the first third of a combustor.

The secondary airflow in the combustion sectionflows at a velocity of several hundred feet per sec-ond around the combustor's periphery. This flow of


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Figure 3-40. As air flows into the combustion section it separates into primary and secondary flows. The primary flow is used tosupport combustion while the secondary flow cools the hot gases before they enter the turbine section.

air forms a cooling air blanket on both sides of theliner and centers the combustion flames so they donot contact the liner. Some secondary air is slowedand metered into the combustor through the

perforations in the liner where it ensurescombustion of any remaining unburned fuel. Finally,secondary air mixes with the burned gases and coolair to provide an even distribution of energy to theturbine nozzle at a temperature that the turbinesection can withstand. [Figure 3-40]

There are currently three basic types of combustionchambers, the multiple-can type, the annular or bas-ket type, and the can-annular type. Functionally,they are the same but their design and constructionis different.

MULTIPLE-CAN TYPE

The multiple-can type combustion chamber con-sists of a series of individual combustor canswhich act as individual burner units. This type ofcombustion chamber is well suited to centrifugalcompressor engines because of the way compressor

discharge air is equally divided at the diffuser.Each can is constructed with a perforated stainless

steel liner inside the outer case. The inner liner ishighly heat resistant and is easily removed forinspection once the combustion can is removedfrom the engine. Each combustion can has a largedegree of curvature which provides a high resis-tance to warpage. However, the shape is inefficientin terms of the amount of space required and theadded weight.

The individual combustors in a typicalmultiple-can combustion chamber are

interconnected with small flame propagation tubes.The combustion starts in the two cans equippedwith igniter plugs, then the flame travels throughthe tubes and ignites the fuel/air mixture in theother cans. Each flame propagation tube isactually a small tube surrounded by a larger tubeor jacket. The small inner tube carries the flame

between the cans and the outer tube carries airflowbetween the cans that cools and insulates. There are8 or 10 cans in a typical multiple-can combustionsection. The cans are numbered clockwise whenfacing the rear of the engine on mostAmerican-built engines, with the number one can

being on the top. All the combustor cans dischargeexhaust gases into an open area at the turbine nozzleinlet. [Figure 3-41]


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Figure 3-41. Used primarily in early turbine engine designs,

multiple-can combustors consisted of a series of individual

burner cans arranged radially around an engine. The multi-

ple-can design has the advantage of easy removal of any

combustor for maintenance or replacement.

ANNULAR TYPE

Today, annular combustors are commonly used inboth small and large engines. The reason for this isthat, from a standpoint of thermal efficiency,weight, and physical size, the annular combustor isthe most efficient. An annular combustion chamberconsists of a housing and perforated inner liner, or

basket. The liner is a single unit that encircles theoutside of the turbine shaft housing. The shroud can

be shaped to contain one or more concentric bas-

kets. An annular combustor with two baskets isknown as a double-annular combustion chamber.

Normally, the ignition source consists of two spark

igniters similar to the type found in multiple-cancombustors.

In a conventional annular combustor, airflow entersat the front and is discharged at the rear with pri-mary and secondary airflow much the same as inthe multiple-can design. However, unlike the cantype combustors, an annular combuster must beremoved as a single unit for repair or replacement.This usually involves complete separation of theengine at a major flange. [Figure 3-42]

Some annular combustors are designed so the air-flow can reverse direction. These reverse-flow com-

bustors serve the same function as the conventionalflow type, except the air flows around the chamberand enters from the rear. This results in the com-

bustion gases flowing in the opposite direction ofthe normal airflow through the engine. This ideawas first employed by Whittle in his early designs.

Figure 3-42. An annular combustor has the highest effi-

ciency for its weight of any combustor design. However,

the engine must be disassembled to repair or replace anannular combustor. Also, the shallow curvature makes this

combustor more susceptible to warping.


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Figure 3-43. A reverse-flow combustor is light and compact.The turbine wheels actually lie within the combustor, ratherthan behind it.

In a typical reverse-flow annular combustor, the tur-bine wheels are inside the combustor area ratherthan downstream, as with the conventional flowdesigns. This allows for a shorter and lighter enginethat uses the hot gases to preheat the compressordischarge air. These factors help make up for theloss of efficiency caused by the gases having toreverse their direction as they pass through the com-

bustor. [Figure 3-43]

CAN-ANNULAR TYPE

Can-annular combustion sections represent a com-bination of the multiple-can combustor and theannular type combustor. The can-annular combus-

tor was invented by Pratt & Whitney and consists ofa removable steel shroud that encircles the entirecombustion section. Inside the shroud, or casing,are multiple burner cans assembled radially aroundthe engine axis with bullet-shaped perforated liners.A fuel nozzle cluster is attached to the forward endof each burner can and pre-swirl vanes are placedaround each fuel nozzle. The pre-swirl vanesenhance the combustion process by promoting athorough mixing of fuel and air and slowing theaxial air velocity in the burner can. Flame propaga-tion tubes connect the individual liners and twoigniter plugs are used for initiating combustion. An

individual can and liner is removed and installed asone unit for maintenance. This design combines the

Figure 3-44. A can-annular combustor contains individual

burner cans in an annular liner. The short burner cans com-bine the compact efficiency of the annular type combustorwith the ease of the multiple-can combustor maintenance.

ease of overhaul and testing of the multiple-canarrangement with the compactness of the annularcombustor. [Figure 3-44]

FLAMEOUT

As mentioned earlier, a combustion flame can beextinguished by high airflow rates. However, exces-

sively slow airflow rates can also contribute to thisproblem. Although flameout is uncommon in mod-ern engines, combustion instability can still occur


25/43


and occasionally causes a complete flameout. Giventhe correct set of circumstances, turbulent weather,high altitude, slow acceleration, and high-speed

maneuvers can induce combustion instability andcause a flameout. There are two types of flameouts,a lean die-out and a rich blow-out. A lean die-outusually occurs at high altitude where low enginespeeds and low fuel pressure form a weak flame thatcan die out in a normal airflow. On the other hand,a rich blow-out typically occurs during rapid engineacceleration when an overly-rich mixture causes thefuel temperature to drop below the combustion tem-

perature or when there is insufficient airflow to sup-port combustion.

TURBINE SECTION

After the fuel/air mixture is burned in thecombus-tor, its energy must be extracted. Aturbine transforms a portion of the kinetic energyin the hot exhaust gases into mechanical energy todrive the compressor and accessories. In a turbojetengine, the turbine absorbs approximately 60 to80% of the total pressure energy from the exhaustgases. The turbine section of a turbojet engine islocated downstream of the combustion section andconsists of four basic elements; a case, a stator, ashroud, and a rotor. [Figure 3-45]

TURBINE STATOR

A stator element is most commonly referred to asthe turbine nozzle; however, you may also hear thestator elements referred to as the turbine guidevanes, or the nozzle diaphragm. The turbine nozzleis located directly aft of the combustion section andimmediately forward of the turbine wheel. Becauseof its location, the turbine nozzle is typicallyexposed to the highest temperatures in a gas tur-

bine engine.

The purpose of the turbine nozzle is to collect thehigh energy airflow from the combustors and directthe flow to strike the turbine rotor at the appropri-ate angle. The vanes of a turbine nozzle are con-

toured and set at such an angle that they form anumber of converging nozzles that convert some ofthe exhaust gases' pressure energy to velocityenergy. In addition, the angle of the stator vanes isset in the direction of turbine wheel rotation. Sincethe gas flow from the nozzle must enter the turbine

blade passageway while it is still rotating, it isessential to aim the gas in the general direction ofturbine rotation. As a result, the velocity energy ofthe exhaust gases is more efficiently converted tomechanical energy by the rotor blades.

CASE

The turbine casing encloses the turbine rotor andstator assembly, giving either direct or indirect sup-

port to the stator elements. A typical case hasflanges on both ends that provide a means of attach-ing the turbine section to the combustion sectionand the exhaust assembly.

Figure 3-45. The four basic elements of a turbine assembly in a

gas turbine engine are the case, stator, shroud, and rotor.

SHROUD

The turbine nozzle assembly consists of an innerand outer shroud that retains and surrounds thenozzle vanes. The number of vanes employedvaries with different types and sizes of engines.The vanes of a turbine nozzle are assembled

between the outer and inner shrouds, or rings, in avariety of ways. Although the actual elements mayvary slightly in their configuration and construc-tion, there is one similarity among all turbine noz-zles: the nozzle vanes must be constructed toallow for thermal expansion. If this is not done,

the rapid temperature changes imposed by theengine would cause severe distortion or warpingof the nozzle assembly.

The thermal expansion of turbine nozzles is dealtwith in several ways. One way requires the vanesto be assembled loosely in the inner and outershrouds. With this method, the shrouds are builtwith a series of contoured slots conforming to theshape of an individual vane. The slots are slightlylarger than the vanes, and therefore, provide aloose fit. In order to provide the strength and

rigidity required, the inner and outer shrouds areencased in an inner and outer support ring. These


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Figure 3-46. A loose fit between the vanes and shrouds

allows thermal expansion to occur without warping the tur-

bine nozzle assembly.

As the high velocity gases pass through the turbinenozzle and impact the turbine blades, the turbinewheel rotates. In some engines, a single turbine

wheel cannot absorb sufficient energy from theexhaust gas to drive the compressor and acces-sories. Therefore, many engines use multiple tur-

bine stages, each stage consisting of a turbine noz-zle and wheel.

The severe centrifugal loads imposed by the highrotational speeds, as well as the elevated operatingtemperatures exert extreme stress on the turbine

blades. At times, these stresses can cause turbineblades to grow in length. If left unchecked, thisgrowth or creep can result in the turbine blades rub-

bing against the engine's outer casing.

support rings also facilitate removal of the nozzlevanes as a unit. Without the support rings, thevanes could fall out as the shrouds were removed.[Figure 3-46]

A second method of attaching the nozzle vanes is torigidly weld or rivet the vanes into the inner andouter shrouds. To allow for expansion, the inner orouter shroud ring is cut into segments. As thermalexpansion takes place, the shrouds expand andclose the gaps between the shroud segments, allow-ing sufficient expansion to prevent stress and warp-

ing. [Figure 3-47]

TURBINE ROTOR

The rotating elements of a turbine section consist ofa shaft and a turbine rotor, or wheel. The turbinewheel is a dynamically balanced unit consisting of

blades attached to a rotating disk. The turbine diskis the anchoring component for the turbine bladesand is bolted or welded to the main shaft. The shaftrotates in bearings that are lubricated by oil

between the outer race and the bearing housing.

This reduces vibration and allows for a slight mis-alignment in the shaft.

TURBINE BLADES

Turbine blades are airfoil shaped componentsdesigned to extract the maximum amount of energyfrom the flow of hot gases. Blades are either forgedor cast, depending on their alloy composition.Early blades were manufactured from steelforg-ings; however, today most turbine bladesconsist of cast nickel-based alloys. In either case,once a blade is forged or cast, it must befinish-ground to the desired shape. As an

alternative to metal turbine blades, thedevelopment of a blade manufactured fromreinforced ceramic material holds promise.Because of ceramic's ability to withstand hightemperatures, greater engine efficiencies may be

possible. Their initial application is likely to be insmall, high speed turbines that operate at veryhigh temperatures.

Turbine blades fit loosely into a turbine disk whenan engine is cold, but expand to fit tightly at normaloperating temperatures. The most commonly used

method for attaching turbine blades is by fir treeslots cut into the turbine disk rim and matchingbases cast or machined into the turbine blade base.[Figure 3-48]

Once installed, a turbine blade may be retained inits groove by peening, welding, rivets, or locktabs.The peening method is used frequently in variousways. A common application of peening requires asmall notch to be ground in the edge of the blade'sfir tree root prior to the blade being installed. Afterthe blade is inserted into the disk, the notch isfilled by the disk metal, which is "flowed" into it

by a small punch mark made in the disk adjacent tothe notch. The tool used for this job is similar to acenter punch. -' -;

Figure 3-47. When vanes are riveted or welded into seg-

mented shrouds, the gaps between shroud segments

allow for thermal expansion.


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Figure 3-48. The loose fit of a fir tree base allows thebase of a turbine blade to expand as it heats to operat-ing temperature.

Turbine blades are generally classified as impulse,

reaction, or a combination impulse-reaction type. Ina turbine that uses impulse blades, the blades merelychange the direction of airflow coming from the tur-

bine nozzle and cause relatively no change in gaspressure or velocity. The turbine wheel simplyabsorbs the force required to change the direction ofairflow and converts it to rotary motion. [Figure 3-49]

Reaction turbine blades, on the other hand, producea turning force based on an aerodynamic action. Todo this, the turbine blades form a series of convergingducts that increase gas velocity and reduce pressure.The result is similar to what happens to an airfoil inthat the reduced pressure produces a lifting force.

Figure 3-49. In an impulse turbine system, the turbine noz-zle vanes form a series of converging ducts that increase

the velocity of the exhaust gases. The impulse turbineblades then extract energy from the gases as the bladesredirect the flow of high velocity gases.

Figure 3-50. The nozzle guide vanes in a reaction turbine directthe exhaust gas flow to strike the turbine blades at a positiveangle of attack. The convergent shape between the turbineblades then increases gas velocity and decreases its pressureto create a component of lift that rotates the turbine wheel.

However, in a turbine, the force is exerted in thedirection of rotation. [Figure 3-50]

To more evenly distribute the workload along thelength of the blade, most modern turbine enginesincorporate impulse-reaction turbine blades. Withthis type of blade, the blade base is impulse shapedwhile the blade tip is reaction shaped. This designcreates a uniform velocity and pressure drop acrossthe entire blade length. [Figure 3-51]

Figure 3-51. To help account for the different rotationalspeeds along the length of a turbine blade, most turbine

engines use impulse-reaction type turbine blades. This typeof blade is constructed with an impulse section at its baseand a reaction section at its tip.


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Figure 3-52. Shrouded blades form a band around the tur-

bine wheel perimeter which helps reduce blade vibration

and increase efficiency.

Turbine blades can be open or shrouded at theirends. Open ended blades are used on high speedturbines, while shrouded blades are commonlyused on turbines having slower rotational speeds.With shrouded blades, a shroud is attached to thetip of each blade. Once installed, the shrouds of the

blades contact each other, thereby providing sup-port. This added support reduces vibration substan-tially. The shrouds also prevent air from escapingover the blade tips, making the turbine more effi-cient. However, because of the added weight,shrouded turbine blades are more susceptible to

blade growth. [Figure 3-52]

To further improve the airflow characteristicsaround shrouded turbine blades, a knife-edge seal ismachined around the outside of the shroud thatreduces air losses at the blade tip. The knife-edgeseal fits with a close tolerance into a shrouded ringmounted in the outer turbine case.

One of the most common ways of cooling the com-ponents in the turbine section is to use engine bleedair. For example, turbine disks absorb heat from hot

gases passing near their rim and from the bladesthrough conduction. Because of this, disk rim tem-

peratures are normally well above the temperatureof the disk portion nearest the shaft. To limit theeffect of these temperature variations, cooling air isdirected over each side of the disk.

To sufficiently cool turbine nozzle vanes and tur-bine blades, compressor bleed air is typicallydirected in through the hollow blades and outthrough holes in the tip, leading edge, and trailingedge. This type of cooling is known as convectioncooling or film cooling. [Figure 3-53]

In addition to drilling holes in a turbine vane orblade, some nozzle vanes are constructed of aporous, high-temperature material. In this case,bleed air is ducted into the vanes and exits throughthe porous material. This type of cooling is knownas transpiration cooling and is only used on sta-tionary nozzle vanes.

Modern engine designs incorporate many combina-tions of air cooling methods that use low and high

pressure air for both internal and surface cooling of

turbine vanes and blades. However, to provide addi-tional cooling, the turbine vane shrouds may also be

perforated with cooling holes.

COUNTER-ROTATING TURBINES

While not common in large engines, some smalltur-boshaft engines feature counter-rotating turbinewheels. Counter-rotating turbines are chosen byengine designers for their effectiveness in dampen-ing gyroscopic effects and reducing engine vibra-tion, not for aerodynamic reasons.

COOLING

When a turbine section is designed, temperature isan important consideration. In fact, the most limit-ing factor in running a gas turbine engine is the tem-

perature of the turbine section. However, the higheran engine raises the temperature of the incoming air,the more power, or thrust an engine can produce.Therefore, the effectiveness of a turbine engine'scooling system plays a big role in engine perfor-mance. In fact, many cooling systems allow the tur-

bine vane and blade components to operate in athermal environment 600 to 800 degrees Fahrenheitabove the temperature limits of their metal alloys.

EXHAUST SECTION

The design of a turbojet engine exhaust sectionexerts tremendous influence on the performance ofan engine. For example, the shape and size of anexhaust section and its components affect the tem-

perature of the air entering the turbine, or turbineinlet temperature, the mass airflow through theengine, and the velocity and pressure of the exhaust

jet. Therefore, an exhaust section determines tosome extent the amount of thrust developed.

A typical exhaust section extends from the rear ofthe turbine section to the point where the exhaust


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Figure 3-53. An internally cooled blade receives cooling air at the root and expels the air at the tip or through holes in the leadingand trailing edges.

gases leave the engine. An exhaust section is com-prised of several components including the exhaustcone, exhaust duct or tailpipe, and exhaust nozzle.[Figure 3-54]

EXHAUST CONE

A typical exhaust cone assembly consists of anouter duct, or shell, an inner cone, or tail cone,three or more radial hollow struts, and a group of tie

Figure 3-54. A typical exhaust section has an exhaust cone, tailpipe, and exhaust nozzle. The exhaust cone is considered the rear-most component of a typical gas turbine engine. The tailpipe and exhaust nozzle are usually classified as airframe components.


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Figure 3-55. The exhaust cone is the rearmost engine com-

ponent. It straightens and smooths the exhaust gas to

extract the greatest possible thrust.

rods that assist the struts in centering the inner conewithin the outer duct. The outer duct is usuallymade of stainless steel and attaches to the rearflange of the turbine case. [Figure 3-55]

The purpose of an exhaust cone assembly is tochannel and collect turbine discharge gases into asingle jet. Due to the diverging passage between theouter duct and inner cone, gas velocity within theexhaust cone decreases slightly while gas pressurerises. Radial struts between the outer shell andinner cone support the inner cone, and help

straighten the swirling exhaust gases that wouldotherwise exit the turbine at an approximate angleof 45 degrees.

TAILPIPE

A tailpipe is an extension of the exhaust section thatdirects exhaust gases safely from the exhaust cone tothe exhaust, or jet nozzle. The use of a tailpipeimposes a penalty on an engine's operating effi-ciency due to heat and duct friction losses. Theselosses cause a drop in the exhaust gas velocity and,hence, the thrust. Tailpipes are used almost exclu-

sively with engines that are installed within an air-craft's fuselage to protect the surrounding airframe.Engines installed in a nacelle or pod, however, often

require no tailpipe, in which case the exhaust nozzleis mounted directly to the exhaust cone assembly.

EXHAUST NOZZLE

An exhaust, or jet nozzle, provides the exhaustgases with a final boost in velocity. An exhaust noz-zle mounts to the rear of a tailpipe, if a tailpipe isrequired, or to the rear flange of the exhaust duct ifno tailpipe is necessary.

Two types of exhaust nozzle designs used on aircraftare the converging design, and theconverging-diverging design. On a convergingexhaust nozzle, the nozzle diameter decreases fromfront to back. This convergent shape produces a

venturi that accelerates the exhaust gases andincreases engine thrust.

The diameter of a converging-diverging ductdecreases, then increases from front to back. Theconverging portion of the exhaust nozzle acceleratesthe turbine exhaust gases to supersonic speed at thenarrowest part of the duct. Once the gases are mov-ing at the speed of sound they are accelerated fur-ther in the nozzle's divergent portion, so the exhaustgases exit the nozzle well above the speed of sound.For additional information on both convergent andconvergent-divergent exhaust nozzles, refer to

Chapter 6, Section B.

On fan or bypass type engines, there are two gasstreams venting to the atmosphere. High tempera-ture gases are discharged by the turbine, while acool airmass is moved rearward by the fan section.In a low by-pass engine, the flow of cool and hot airare combined in a mixer unit that ensures mixing ofthe two streams prior to exiting the engine. High

bypass engines, on the other hand, usually exhaustthe two streams separately through two sets of noz-zles arranged coaxially around the exhaust nozzle.

However, on some high bypass engines, a commonor integrated nozzle is sometimes used to partiallymix the hot and cold gases prior to their ejection.[Figure 3-56]

An exhaust nozzle opening can have either a fixedor variable area. A variable geometry nozzle issometimes necessary on engines that utilize anafterburner. Variable nozzles are typically operatedwith pneumatic, hydraulic, or electric controls.

AFTERBURNERS

Afterburners are used to accelerate the exhaust gases,which in turn, increases thrust. An afterburner is


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Figure 3-56. On some high bypass engines, cold bypass air mixes with hot exhaust gases after the gases exit the engine. Other

high bypass engines use a common or integrated exhaust nozzle that partially mixes the gas streams internally.

typically installed immediately aft of the last stageturbine and forward of the exhaust nozzle. The com-

ponents that make up an afterburner include the fuelmanifold, an ignition source, and a flame holder.[Figure 3-57]

The addition of an afterburner to a gas turbineengine is made possible by the fact that the gases inthe tailpipe still contain a large quantity of oxygen.If you recall, approximately 25 percent of a com-

pressor's discharge air is used to support combus-tion, while the remaining 75 percent is used forcooling. Once the cooling air passes through anengine, a portion of it is mixed with the exhaustgases at the rear of the turbine section. The tailpipeentrance is fitted with a fuel manifold, consisting ofa set of afterburner fuel nozzles, or spray-bars, thatinject fuel into the tailpipe. The fuel and air mix,

then ignite and burn in the afterburner. The addi-tional heat generated by combustion accelerates theexhaust gases and creates additional thrust.

Figure 3-57. An afterburner is used to increase thrust andconsists of a fuel manifold, an ignition source, and a flame

holder.


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Figure 3-61. By referring to the graph above, you can see

that the decibel level of early turbojet aircraft without noise

suppression equipment typically exceeded 110 decibels. In

comparison, modern high bypass turbofan engines produce

less than 100 decibels of sound.

Turbojet engines, especially older designs,frequently require additional noise suppressionequipment. The additional equipment typicallyincludes a device that breaks up airflow behind thetail cone, and some new forms of sound insulatingmaterial. In addition, some airframe components

are being redesigned and installed as noisereduction kits to meet new Federal standards.

The sound intensity of engine noise levels ismeasured in decibels (db). A decibel is the ratio ofone sound to another. One db is the smallestchange in sound intensity that the human ear candetect. Approximately 60 db is a comfortable levelfor conversation and background music. On theother hand, loud music peaks at more than 100 db.Sound that reaches approximately 130 db can cause

physical pain. [Figure 3-61]

To help answer the concerns over noise aroundairports, the Federal Aviation Administration hasestablished guidelines for aircraft operators thatspecify maximum noise limits based on aircraftweight. All older aircraft that utilize louder turbojetengines were given a grace period to modify theengines to meet the maximum noise levels. [Figure3-62]

ENGINE MOUNTS

Engine mount design and construction for gasturbine engines is relatively simple. Since gasturbine engines produce little torque, they do notneed heav-

Figure 3-62. The curves on the two graphs shown here rep-

resent the maximum decibel levels aircraft are allowed to

produce during the takeoff and approach phases of flight.

Below each curve are the decibel levels produced by differ-

ent aircraft.

ily constructed mounts. The mounts do, however,support the engine weight and allow for transfer ofstresses created by the engine to the aircraftstructure. On a typical wing mounted turbofanengine, the engine is attached to the aircraft by twoto four mounting brackets. However, because ofinduced propeller loads, a turboprop developshigher torque loads, so engine mounts are

proportionally heavier. By the same token,turboshaft engines used in helicopters are equipped

with stronger and more numerous mount locations.[Figure 3-63]

BEARINGS

The combination of compressor and turbine rotorson a common shaft make up the main engine power,or rotor shaft, which must be adequately supported.Engine main bearings are assigned that criticalfunction of support and are located along thelength of the rotor shaft. The number of bearingsnecessary is determined, in part, by the length andweight of the rotor shaft. For example, since asplit-spool axial compressor typically has a greaternumber of rotating components it requires moremain bearings than a centrifugal compressor.


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Figure 3-65. Propellers driven by a free turbine rotate independently of the compressor turbine.

downstream from the main turbine and is dedicatedto driving only the propeller. [Figure 3-65]

A second method of transferring the exhaust gasenergy to the propeller is through a fixed shaft. In

this case, the main turbine typically has anadditional turbine wheel that extracts the energyneeded to drive a propeller. With these fixed shaftengines, the main power shaft goes directly into areduction gearbox to convert the high-speed lowtorque turbine output into low-speed high torqueenergy to drive a propeller.

To prevent the propeller on a turboprop engine fromdriving the turbine, a sophisticated propellercontrol system must be used to adjust propeller

pitch as necessary to match the engine's output. For

example, in normal cruise flight, both the propellerand engine rpm remain constant. Therefore, inorder to maintain a constant-speed condition, the

propeller's blade angle and fuel flow must beadjusted simultaneously. In other words, whenfuel flow is increased, the propeller's pitch mustalso increase.

TURBOSHAFT ENGINES

Turboshaft engines are gas turbine engines thatoperate something other than a propeller by deliv-ering power to a shaft. Turboshaft engines aresimilar to turboprop engines, and in someinstances, both use the same design. Liketurboprops, turboshaft engines use almost all theenergy in the

exhaust gases to drive an output shaft. The powermay be taken directly from the engine turbine, orthe shaft may be driven by its own free turbine. Likefree turbines in turboprop engines, a free turbi

3 Turbine Engines (1)saf

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