koff gas turbine technology evolution a designers perspective

8/10/2019 Koff Gas Turbine Technology Evolution a Designers Perspective

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American Institute of Aeronautics and Astronautics1

GAS TURBINE TECHNOLOGY EVOLUTION - A DESIGNERS PERSPECTIVE

Bernard L. Koff, AIAA FellowTurboVision, Inc.

Palm Beach Gardens, Florida

Abstract:

During the past 50 years the aircraft gas turbine has evolved into the worlds most complex product whichhas made an astoundingly positive impact on mankind. Jet powered aircraft have provided the UnitedStates with unprecedented air power supremacy for defense and global reach to help promote worldwidepeace and aid. Large turbofan powered transport and commercial aircraft have spanned the Globe,making the world much smaller while clean burning gas turbines are used worldwide for powergeneration. Lessons learned and design innovations developed for gas turbines have also beentransitioned to rocket engines including the oxygen and hydrogen pumps for the space shuttle mainengines. This presentation highlights key technologies created and developed by engineers and whichhave been responsible for the extraordinary evolution of state-of -art advances in gas turbine propulsion.

In the Beginning

The Americans got a late start in the developmentof the gas turbine engine because responsibleleaders didnt believe the gas generator cycle

consisting of a compressor, combustor and turbinewas practical. The reasoning was that after thepower was extracted from the turbine to drive thecompressor, there wouldnt be enough residualenergy in the exhaust gas for useful work. Thisreasoning was partially dispelled in the U.S. afterthe first flight of the Gloster aircraft in 1941powered by the Whittle jet engine. Whittle was avisionary genius with excellent design engineeringskills and great determination. His 1930 enginepatent (Figure 1) shows a compressor with twoaxial stages followed by a centrifugal stage, anaxial cannular combustor with fuel nozzles and a

two stage axial turbine.

Figure 1 Whittle Gas Turbine DrawingBritish Patent No. 347206

(Courtesy of Rolls Royce)

It would be an understatement to say that Whittlehad great difficulty in getting support to pursue hisrevolutionary invention. However, after persistingwith great courage and personal sacrifice he wasable to successfully test the kerosene fueled W.1in 1937, the Worlds first jet engine.

Four years later, General Hap Arnold was briefedon the progress of the Whittle engine and theGloster aircraft shortly before the first flight and

began working to obtain production rights for theUnited States. The improved W2 engine model(Figure 2) incorporated a double sided centrifugalcompressor, an axial reversed flow cannularcombustor, a single stage axial turbine andproduced a thrust of 1560 lb. The exhaust gasenergy or specific power reached approximately50 hp for each lb/sec of airflow.

After negotiations with the British government,Frank Whittle was sent to the United States toteach the Americans how to design and build hislatest W.2B jet engine.

Figure 2 Whittle W2 Engine1

The General Electric Company was selected sincethey had already designed and manufacturedturbo-superchargers for reciprocating engines.This effort resulted in producing the GE I-A, thefirst American jet engine which was a copy of theWhittle W.2B engine. Two I-A engines poweredthe Bell XP-59A aircraft to Americas first jet flightin late 1942.

AIAA/ICAS International Air and Space Symposium and Exposition: The Next 100 Y4-17 July 2003, Dayton, Ohio

AIAA 2003-272

Copyright 2003 by Bernard L. Koff . Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


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Unknown to the Allies, Hans von Ohain, a brilliantengineering student, with continuing support fromErnst Heinkel the aircraft manufacturer, haddeveloped a petrol fueled jet engine by March1938, shortly after Whittle. In von Ohains ownwords, Heinkel was crazy for speed and gave useverything. This resulted in the first jet poweredflight in the summer of 1939 on the eve of WorldWar II. The von Ohain engine (Figure 3) had anaxial flow inducer ahead of the centrifugal impellerstage, a reverse flow combustor and a radialinflow turbine. The exhaust gas energy wasapproximately 50 hp for each lb/sec of inlet airflow,similar to Whittle.

Figure 3 Hans von Ohains jet engine He S 3B2

During World War II, the Junkers Jumo enginepowering the famous Messerschmitt Me 262 jet

aircraft was developed in a competition withHeinkel. The Jumo engines (Figure 4) for thisaircraft were mounted in nacelles rather thaninternal to the aircraft fuselage using an axial flowcompressor, axial flow turbine and straight throughflow combustor to reduce frontal area andincrease performance. This early configurationbecame a forerunner of how future jet engineswould be configured relative to overall designarrangement.

Figure 4 Junkers Jumo 004 Turbojet3

The inventor designers Whittle and von Ohaintalked about past achievements at the symposiumcommemorating von Ohains 50th anniversary of

the first jet flight (Figure 5). They suggestedincluding me who talked about the future.

Figure 5 Speakers von Ohain, Koff and Whittle50th Anniversary of first flight

(Dayton Engineers Club 1989)

Advancing the Technology

Materials

Identifying key technologies over the past 50 yearsresponsible for the evolution of the gas turbineappropriately begins with the achievements of thematerials and manufacturing process engineers. Achronological progress (Figure 6)of turbine airfoilmaterial capability over the past 50 years showsan improvement exceeding 500

oF. This

achievement was the result of many innovativescientific breakthroughs in materials research,processing and manufacturing technology.

Figure 6 Turbine Airfoil Materials Progress

The early jet engines were severely limited by hotsection materials motivating the research andinvention of improved alloys. The concept ofdeveloping superalloys was discovered in the 40sbut the air melting process produced low ductilitywith the addition of the key strengtheningelements such as aluminum and titanium. The


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breakthrough came in 1953 when VacuumInduction Melting (VIM) was developed, a processheralded as having made the jet engine what is ittoday. This innovative process boosted alloycapability by 200

oF for turbine airfoils in the 1955

time period (Figure 6).The Vacuum Arc Remelting(VAR) process followed in 1958 to produce largeforgings for disks. These innovative manufacturingprocesses launched the development of todaysgeneration for disks, shafts, bolts and structures.

An interesting observation is that if an alloy can beused to produce a bolt requiring an upset forgedhead for strength, it can also be used for mostother engine components such as compressorairfoils, disks, casings and frames.

The combustor hot streak gas entering the turbinemade it necessary to air cool the stationary vaneson the early engines and X-40 castings were astandard for some 30 years. Since the rotor blades

pass through the hot streaks, they experience alower average temperature. Also, since the rotorblades are moving they are subjected to a lowerrelative temperature. This allowed the use offorged alloy blades with both higher fatiguestrength and mechanical properties for more than25 years. The introduction of air cooled bladesrequired complex internal passages more readilyprovided by lower strength castings but with alsohigher temperature capability. Dampers under theplatforms of the cast blades also became standardfeatures to suppress vibratory response. HotIsostatic Pressing (HIP) (Figure 7)was introduced

in the early 70s to reduce porosity and increaseboth ductility and fatigue strength for castings.

Figure 7 Hot Isostatic Pressing (HIP) Process

The introduction of Directional Solidification (DS)and Single Crystal (SC) superalloys (Figure 8)produced a breakthrough providing a 200

oF

increase in metal temperature capability overconventional multigrain equiaxed cast materials.

Figure 8 Turbine Airfoil Material Evolution

Equiaxed castings have many grain boundariessurrounding the crystals of the superalloy formingfailure initiation points in fatigue, creep andoxidation. The DS castings arrange the crystals inthe form of radial stalks eliminating the weakergrain boundaries in the tensile direction providing

improved resistance to thermal fatigue and creep.The SC casting process goes one step further bycompletely eliminating all weaker grain boundariesand providing further improvements in resistanceto creep, fatigue and oxidation. This highlyinnovative process originally invented by amaterials research engineer has made it possibleto cast a complete turbine airfoil, dovetail andplatform in a single superalloy crystal.

An additional benefit of the DS and SC alloys isthat they can be tailored in the casting process todirectionally exhibit a lower Youngs modulusresulting in a lower stress for the same strainrange. This feature allows the designer to pass thelowest temperature cooling air up through theinternal blade passage at the hot airfoil leadingedge. Higher airfoil cooling efficiency can then beachieved without encountering thermal fatiguecracks experienced in equiaxed cast blades.

Realizing that nickel based superalloys encounterincipient melting at 2400

oF, work began to

develop thermal barrier coatings for hot sectionairfoils to prevent oxidation. While superalloydevelopment proceeded, Aluminide coatings werefirst applied in the mid 70s to meet the demand for

increased hot section life. Ceramic thermal barriercoatings (Figure 9) were applied in the mid 80safter reaching within 400

oF of incipient melting

(shown in red) with the best directionally solidified(DS) and single crystal (SC) alloys.Thermal barrier coatings are also applied to theblade outer air seals (BOAS) which are subjectedto higher combustor temperatures than the blades.


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Figure 9 Thermal Barrier Coatings

Turbine Airfoil Cooling

Today, the jet engine turbine blade is the Worldsmost sophisticated heat exchanger. Until the mid60s, there were three basic schools of thought for

the design of the first stage high pressure turbineblade; uncooled, convection cooled and filmcooled. Many advocated that putting cooling holesin the highly stressed turbine blades would lead tofailures. Others demonstrated in the late 50s thatconvection cooled blades using radial holes drilledinto the core of the airfoil with a Shaped TubeElectrolytic Machining (STEM) process would notcompromise fatigue strength. Material removalusing the STEM process did not leave a brittlerecast layer subject to cracking.

The film only group maintained that a film of coolair should be used as a barrier between the hot

gas and metal. Turbine blades were designed andmanufactured in the late 50s using a forged radialstrut and dovetail with a brazed on porous sheathforming the airfoil. The concept of the poroussheath airfoil was to discharge air on the airfoilsurface to achieve transpiration cooling whileprotecting the load carrying strut. This conceptwas not successful because of backflow when thehot gas in flowed and mixed with the airfoil internalcooling air.

Its interesting that for many years, both cast andfabricated turbine stator vanes were successfullyusing film holes on the airfoil leading edge to cool

the combustion hot streaks. Why the film cooledvane leading edge technology was not adapted toblades has been a contentious issue. A likelyreason for not considering leading edge film holesto cool the rotating blades was the fear ofencountering high cycle fatigue failures.

A breakthrough was made in the early 70s whenthe Air Force funded industry to develop a turbineblade using both convection and film cooling. The

engineers worked closely with casting suppliers toproduce a one piece casting of a convection/filmcooled blade as well as methods for producingboth round and shaped film cooling holes.

Finally, in the early 80s, the design, materials andprocessing came together to produce a one piece

convection/film cooled blade using single crystal(SC) materials (Figure 10).

Figure 10 Single Crystal Turbine Blade with FilmAnd Convection Serpentine Cooling

Internal cored passages form the compartmentswhere compressor cooling air flows in a five passserpentine and a single passage flowing air to acavity adjacent to the leading edge. The internalpassages have cast in trip strips to promoteturbulent flow and increase the convection heattransfer coefficient. Film cooling is provided bystrategically discharging air on the airfoil concavepressure surface. The film cooling holes at the

leading edge are densely spaced to provide bothconvection and film cooling where the hot gas heattransfer rate is highest. The suction (convex) sideof the airfoil has shaped cooling holes to help keepthe film attached to the surface. All suction surfacecooling holes are also upstream of the airfoilpassage throat to minimize mixing losses. Thetrailing edge has pin-fin pressure side dischargecooling to minimize thickness and reduce thewake loss. A milestone was achieved in the early80s when the blade shown successfully passedan accelerated 4000 Tactical Air Command (TAC)cyclic endurance test involving rapid hot starts andthrottle retards.

In the 60s, the Tangential On-Board Injector(TOBI) concept (Figure 11) was invented by anengineer at P&W to lower the temperature of thecompressor discharge cooling air before it enteredthe first stage turbine blade. The conceptdeveloped into an annulus with turning vanes toaccelerate the airflow from axial to tangential rotorspeed decreasing the temperature whileminimizing the pressure loss entering the rotor.


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The pressure drop across the combustor and firststage turbine vane accommodates the TOBIpressure drop and also allows the dischargepressure to be set higher than the turbine flowpathto avoid backflow into the airfoil.

Cooling

flow

Turbine

blade

TOBI

Turbine

disk

Stator

Compressor

Figure 11 PW F100-229 Turbine Rotor TOBICooling the Cooling Air to the Blade(Courtesy of P&W)

The TOBI decreases the temperature of thecompressor cooling air by 125

oF for the 2 stage

turbine shown. The TOBI also decreases theturbine pump work required in getting the air up torotor speed before entering the blade. Engineswith cooled turbine blades and interstage vanecavities have used the TOBI concept for the past25 years to reduce the temperature of the coolingair, increase efficiency and improve durability.

The turbine vanes must accommodate combustor

hot streaks, depending on the pattern factor, whichcan be 400

oF higher than the average gas

temperature. Since the vanes are stationary, theyare subjected to the total gas stagnationtemperature. The turbine vanes (Figure 12) usedin the 4000 TAC cycle accelerated endurance testhave extensive internal convection and externalfilm cooling to meet the higher gas temperatures.These single crystal film/convection cooled vanesset a milestone in durability for high temperaturegas turbine engines.

Typically, the stationary turbine vanes for aircraftengines require approximately 10% of the inletcompressor flow for cooling with combustordischarge temperatures in the range of 2800-3200

oF. At the average gas temperature, the

rotating turbine blades typically use 4% of thecompressor flow for cooling. Although the turbineblades operate at lower gas temperatures than thevanes, the metal temperatures must be reduced toaccount for centrifugal and vibratory stresses.

Figure 12 Single Crystal Turbine Vanes Using FilmAnd Convection Cooling

A chronological evolution of higher turbine rotorinlet temperature (RIT) capability as a function of

the cooling effectiveness using single crystalmaterials is represented in Figure 13.

Figure 13 Turbine Blade Cooling TechnologyInlet Gas Temperature vs. Effectiveness

The cooling effectiveness is represented by theratio of the airfoil heat load to cooling flow, ameasure of how well the airfoil is cooled betweenthe hot gas and cooling air temperatures. The RITbase for the solid uncooled blade is 1800

oF

representing the mid-50s technology. Note thatwithout cooling, there is only a 50

oF improvement

in RIT in going from the first to the latestgeneration of single crystal material. Progressingto convection cooled blades with a coolingeffectiveness of 0.4 allows a 400

oF increase in

RIT. The payoff for increased material temperaturecapability is amplified as the cooling effectivenesslevel increases. The single crystal film/convectioncooled blade family with an effectiveness of 0.6+has RIT capability to 3000

oF. This operating

temperature is 1200oF above the solid uncooled


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blades and 600oF above the 2400

oF incipient

melting temperature of nickel based superalloys.This spectacular progress resulted from adedicated team effort that combined single crystalmaterial manufacturing processes, innovativecasting suppliers, creative designers andgovernment support.

Compressor Design and Engine Configuration

The compressor has often been referred to as theheart of the engine. Air must be pumped todischarge pressure at high efficiency withoutencountering failures or stall instability induced byinlet distortion, Reynolds Number effects, enginetransients, acceleration back pressure and controltolerances. The normal compressor operating lineand stall line (a compression limit) is shown as afunction of pressure ratio and airflow (Figure 14).

Figure 14 Compression System Stability Audit

Lessons learned have repeatedly demonstrated

that its essential for the engine compressor toaccommodate the factors shown in the stabilityaudit that have the potential for causing a flowbreakdown. Inlet distortion caused by flowseparation and also Reynolds effects due to thelower air density at altitude decreases the stallline. Control tolerances and acceleration fuel flowthat increases combustion back pressure bothraise the operating line. Transient thermals,deterioration and hardware tolerances drop thestall line and raise the operating line. The key is tohave enough compressor stall margin remaining(SMR) for safe engine operation. Higher rotor tip

speeds, low aspect ratio airfoils and axial inletvelocity to wheel speed ratios (Cx/U) in the rangeof 0.4-0.5, have substantially raised stall margin.

The early engines were all single rotor turbojetswith fixed geometry compressors. In the late 40s,GE developed the J47 turbojet (Figure 15) with apressure ratio of 5 and a Curvic coupling rotor toreduce engine vibration resulting from shiftingparts experienced on the earlier J35. The turbine

rotor disk had a Timkin 1625 alloy rim TIG weldedto an AMS 4340 hub to provide higher strength atthe flowpath. Heating of the Aluminum inlet guidevane struts with compressor bleed was laterdeveloped and added to prevent ice buildup.

Figure 15 GE J47 Single Rotor Turbojet Engine(Courtesy of GE)

By the early 50s, two schools of though haddeveloped. In a major step forward with the goal toleapfrog the industry, Pratt & Whitney successfully

developed the J57 (Figure 16), an axial flow dualspool turbojet with a 9 stage LP compressor, anda 7 stage HP compressor driven by a single stageHP and 2 stage LP turbine. Thrust was 10,500with an initial pressure ratio of 11 and a thrust toweight of 2.7.

Figure 15 P&W J57 Dual Spool Turbojet4

Meanwhile, GE developed the single rotor J79turbojet (Figure 17), with a variable geometry 17stage compressor at a pressure ratio of 13.3 anddriven by a 3 stage turbine.

Figure 17 GE J79 Turbojet with Afterburner5


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A hydromechanical control scheduled the variableinlet guide vanes and front 6 stators to match frontand rear stages during acceleration.

P&W argued that having two compressors onseparate shafts with bleed for matching stages,would operate closer to their optimum corrected

speeds producing higher pressure ratios andoperational flexibility. GE insisted that the J79single rotor turbojet with variable stators was lesscomplex with fewer parts and lower manufacturingcost. For a considerable time, both companieswere polarized in their views. Eventually and withcompelling reasons, engine configurations usingdual spool, bleed matching and variable statorswere adopted by both companies.For subsonic flight, the exhaust velocity of theturbojet engine is significantly higher than theaircraft flight speed reducing propulsion efficiencyand increasing fuel efficiency.

Rolls Royce is credited with developing theconcept of first using fan stages to bypass airaround the core engine to reduce the exhaust jetvelocity for improved subsonic performance. Theinitial reaction of P&W and GE was negative andseemingly confirmed when the Rolls RoyceConway turbofan engine didnt outperform the

turbojet. It has been speculated that the fan had tohave low efficiency to produce such a result.

In an effort to enter the commercial aircraft enginemarket, GE added an aft fan module to the J79turbojet and this became the first Americanturbofan engine. The GE aft fan rotor consisted ofa turbine blade supporting a tip mounted fan bladeseparated by a transition platform and seals. Thetemperature gradient across the platform betweenthe tip of the turbine blade and root of the fanblade was 400

oF requiring considerable

development to eliminate low cycle fatigue.

P&W countered the GE aft fan with the successful

JT3D/TF33 front fan engine and succeeded incapturing the market on the B52, B707, DC-8 andmany other aircraft. It soon became evident thatP&Ws front fan configuration was superior since itsupercharged the core compressor and alsoproduced lower nacelle drag.

The mechanical design of the multi-stage axialcompressors proved to be a formidable problemfor the engineers from the beginning since:

the airfoils have relatively thin edges forperformance and are subject to damage

high aspect ratio airfoils have been easierto design for high efficiency but have lowerstall margin and durability

first stage blades can encounter flutter

vibration at low corrected speed (highMach) caused by high angle of attack

blade dovetails must be stronger than theairfoils and disk dovetails even stronger

rotor disk and shaft assemblies must notchange balance after assembly and

close operating clearances are requiredfor high efficiency and stall margin.

The variable vanes must also track accurately toprevent stall and blade fatigue failures.

Supersonic aircraft engines began encounteringfirst stage blade flutter (self excited vibration) atthe higher flight Mach Numbers. Airfoil torsional

flutter can occur when the compressor operates atlow corrected speed due to the high ram inlettemperature. In the late 50s, the GE YJ93 turbojetengine encountered first stage blade failures atMach 2.2 and used loose fitting pins from blade toblade to damp the flutter vibration. The engineersrealized that another solution was needed for thePFRT (Preliminary Flight Rating Test). Increasedchord to add beam stiffness was rejected becauseof the excessive weight increase. The break camewhen a GE design engineer saw a scrapped P&WJT3D fan blade with a midspan shroud at a vendorsite. J93 blades were then manufactured using theP&W midspan shroud idea to pass PFRT. TheJT3D midspan shroud blade design is still usedthroughout the world for moderate aspect ratio fanand compressor blades (Figure 18).

Figure 18 PW4000 Midspan Shrouded Fan BladeThe shroud is formed by angel wing extensionsintegrally forged with the blade which butt togethermidway in the flowpath providing stiffness againstflutter. The contact areas are coated with tungstencarbide to resist fretting and wear. The original midspan shrouds had flat plate cross sections. Astreamlined airfoil cross section for the midspanshroud was later adopted to reduce shroud dragand efficiency loss.


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Newer turbofan engines incorporate low aspectratio and low radius ratio fan blades withoutshrouds to improve efficiency, stall margin and theratio of flow per unit annulus area. The largest ofthe commercial turbofan engines use both hollowdiffusion bonded titanium and composite blades toreduce weight with the low aspect ratio airfoils.

Over the years, aircraft engine rotor configurationvaried with company experience and included:

Curvic Coupling teeth machined intointegral disk spacers with tie bolts toclamp the assembly

Spacers bolted to disks using rabbets (topilot) or close fitting dowel bolts for axialclamping and radial positioning

TIG (tungsten inert gas), Plasma Arc andElectron Beam welding to attach diskspacers

Inertia welding to attach disk and spacers

(developed at GE in 1968)Over a 15 year period, the superiority of inertiawelding for compressor and turbine rotors wasestablished and implemented (Figures 19 and 20).These rotors provided maximum rigidity andbalance retention while insuring low maintenance.

Figure 19 GE F110 Inertia Welded CompressorRotor in Titanium & Inconel 718 Superalloy

Figure 20 PW F100 Inertia Welded CompressorRotor Spool and Shaft in Inconel 718

Inertia welding is a forging process that caneliminate defects and achieve parent metalstrength. Energy is stored in a flywheel where a

rotor part is mounted and then moved into contactwith a stationary part. Forging of the rotor takesplace as the flywheel energy is dissipated. Whenthe parameters are set properly, there is nomelting and resolidification to produce defects.

Metal upset or weld flash on both spacer surfaces

should be machined and shot peened for surfaceenhancement. Except where the titanium stagesare attached to the higher temperature nickelalloys, welded rotor spools eliminate bolt holesand stress concentrations in the disk web and rim.

Aircraft core engine compressors (without lowspool supercharging) have increased in pressureratio from 2 to 20 while efficiencies increased from78 to 90%. Typical fighter engine compressors(Figure 21)have pressure ratios of 8 while somelarge commercial engines such as the GE90, havecompressor pressure ratios over 20.

Figure 20 PW F100-229 Core Engine CompressorInertia Welded Rotor in Titanium & IN718

The PW F00-229 core compressor rotor is inertiawelded with only one bolted joint between theforward 2 stage titanium spool and a 7 stageIN718 spool. The variable inlet guide vanes arefollowed by 3 variable vane stages. The internalrotor drum is vented and cooled by third stage airto improve the thermal match with the outer casingfor clearance control. This concept, firstimplemented at GE in the mid 60s, is now used

worldwide for cooling and reducing the rotor andstator radial clearances for improved performance.

The GE90 commercial core engine compressorhas the Worlds highest pressure ratio at morethan 2.5 times the military engine for the samenumber of stages. This is a tribute to theengineers who worked for many years to increasethe average stage pressure rise without incurring aserious efficiency penalty (Figure 22).


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Figure 22 GE90 Core Engine Compressor(Courtesy of GE)

As pressure ratios increased, radial clearances forblades and vanes became increasingly important.Since metal to metal rubs can induce failure, bothabradable and abrasive coatings were developedand put into service during the past 40 years. Theabradable coatings were rough causing aperformance loss and also spalled leaving cratersin the blade tip shrouds which reduced stallmargin. The abrasive coatings loaded up withmetal debris during a rub causing a metal to metalrub with local overheating.

In the mid 80s, Cubic Boron Nitride (CBN) bladetip coatings for compressors and turbines weredeveloped to prevent blade tip wear during lightrubbing. The application of CBN to the blade tips

increases initial cost but allows radial clearancesto be reduced for increased efficiency withoutencountering rub damage (Figure 23).

Figure 23 Compressor Blade Tip CoatingCubic Boron Nitride (CBN) Grits

The CBN grits improved the compressorefficiency by 1% by allowing the blades to rub intothe stator shroud in the range of 10-15 mils.Without CBN coatings, the radial clearance atassembly would have to be increased to insure asafe margin against rubs. Newer turbomachinerycomponents are using CBN coatings successfully.

In the early engines, it was common for interstageknife edge or labyrinth seals between the rotor andstator vanes or casing structure to run opposite asolid metal seat. It took considerable experience toset a radial gap that would avoid a rub that couldmelt metal and also minimize leakage. To avoidexcessive heating during a rub, many labyrinthseal teeth were machined with a thin ribbon ofmaterial at the tip. After numerous failures, a GEengineer invented a honeycomb seal seat in the50s that could be rubbed without causing distressto the labyrinth teeth (Figure 24). It took time forthe engineers to gain confidence in setting thelabyrinth teeth to rub but this configuration wasuniversally adopted by all engine manufacturers.

Honeycomb

LabyrinthTeeth

Rotor spacer

Stator shroud

Figure 24 Honeycomb Labyrinth Seal

In the 60s, the engineers began to coat therotating labyrinth teeth with Al2O3 in criticallocations to prevent excessive wear during initialbreak-in and extreme transient operation. Coatedteeth have become standard design practice toimprove durability and reduce maintenance cost.

The J79, J93, TF39 and CF6 family of enginecompressors used a smooth spool configurationthat eliminated interstage seals by having the vanetips run close to the rotor shell. As late as the early70s some engineers argued that a smooth spoolrotor design gave superior performance over onewith interstage shrouds even with honeycomblabyrinth seals. However, in a back to back test onthe GE F404 engine compressor, it wasestablished that stator vane honeycomb shroudsand rub in labyrinth seals resulted in higherperformance. As the aerodynamic stage loadingincreased, it became evident that shrouded stator

vanes were more durable than cantilever vanesand provided higher resistance to vibratory modes.Efforts to improve compressor performance andoperability highlighted the need to improve controlof radial clearances. Unequal thermal heating andcooling of the rotor and stator for compressors andturbines result in increased radial build-upclearances to avoid excessive rubbing duringtransients. For the compressor, a combination ofcirculating cooler air within the internal rotor


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cavities and slowing the stator thermal responseusing shielding and lower expansion alloys hasreduced thermal clearance excursions (Figure 25).

Figure 25 Matching Compressor Rotor and StatorRadial Clearances to Reduce Leakage

In the case of the turbine, an active clearancecontrol concept (Figure 26) was developed by

P&W using compressor cooling air on internalparts adjacent to the flowpath and coolermodulated fan air on the turbine outer casing

Figure 26 Active Clearance Control to ImproveThermal Match and Reduce Leakage

The introduction of Computational Fluid Dynamics(CFD) (Figure 27)allowed the aerodynamicists tomodel the complex flow field in three dimensions.

Figure 27 Bowed Compressor Stators ProvideHigher Efficiency and Stall Margin

CFD-3D modeling for steady and unsteady flowprovides a more complete physical understandingof the airfoil flow field from root to tip and hasreplaced many empirical correlations with physics.The CFD modeling has also been very useful inthe understanding and solution of complexstructural dynamics problems involving turbulentand separated flow.

The average compressor discharge temperature(T3) has been limited to approximately 1200

oF for

the past 40 years limiting the maximum pressureratio and flight speed (Figure 28). At 1200

oF,

nickel superalloys used in the high speed rotatingdisks can encounter creep and rupture. At sealevel takeoff Mach 0.25, pressure ratios of 40-42reach this T3 limit. At Mach 3 flight speed in thealtitude range between 36,000 and 69,000 ft, theair is already compressed to 632

oF at the

compressor inlet. This high inlet temperature limits

engine pressure ratios to a range of 3.5-3.8without exceeding the peak profile temperaturelimit at the compressor discharge. For the future,higher temperature superalloys are needed toextend pressure ratio, flight speed andperformance of air breathing turbine engines.

Figure 28 Compressor Exit Temperature LimitsPressure Ratio and Flight Mach number

Commercial aviation has spanned the globeusing evolutionary advances in jet propulsion.Increasing compressor pressure ratios has been akey in achieving higher engine cycle efficiency forlonger range aircraft. Pressure ratios for subsonicaircraft applications have increased by a factor of

20 during the past 50 years (Figure 29). The morerecent engines have reached pressure ratios of 40at sea level and are already pushing thecompressor discharge temperature limit withcurrent nickel superalloys.


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Figure 29 Compressor Growth in Pressure Ratio

Combustion Design

In the late 50s efforts began to eliminate visiblesmoke particulates and by 1970, work was inprogress to reduce unburned hydrocarbons (HC)and carbon monoxide (CO). Combustion efficiency

at high power has always been near 100% butduring the next ten years, it was improveddramatically at or near idle power. Additionalprogress was made in the 80s in understandingand developing technologies to reduce oxides ofnitrogen (NOx) emissions. With gas turbinecombustion efficiencies above 99%, littleopportunity existed for further reductions in HCand CO. Smoke was reduced below the thresholdof visibility with lean combustion while stillproviding adequate windmilling air starts. Thefocus in the 90s has been on NOxemissions dueto its contribution to ground level ozone and smog,

acid rain and atmospheric ozone depletion.

Reduced emissions combustors have also beensignificantly reduced in axial length, fuel deliverynozzles have been improved to eliminate fuelcoking and overall durability has been increasedby orders of magnitude (Figure 30).

Figure 30 Combustor Durability Improvements

Shorter combustors reduce the engine axial lengthand rotor bearing span, require less liner coolingair and improve the combustion pattern factor.

The durability of the combustion liners enclosingthe hot gases became an issue with increasedengine life requirements. For many years,combustion liners were manufactured using spotand seam welded overlapping sheet metal louverswith relatively short life (Figure 31).

Figure 31 Increasing Combustor Liner Durability

The machined ring liners provided a 10X life

improvement over the sheet metal liners at thesame combustion temperatures. In the early 70s,the Air Force Materials Laboratory funded a newconcept which provided a liner panel shielding theouter liner shell from the hot gases. This conceptwas developed resulting in a 10X life improvementover machined ring configurations in use and a100X life improvement over the earlier sheet metalliners. Improved machined ring liners have alsobeen developed using film/convection cooling.

High Mach Engines

In the mid 50s there was considerable Air Forceand Navy interest in high mach flight as the nextfrontier for reconnaissance, interceptor andbomber aircraft. GE was funded to develop theYJ93 engine (Figure 32) for the XB-70 Mach 3supersonic bomber. Two prototype aircraft werebuilt and flight tested before the program wascancelled. The original compressor had 12 highaspect ratio stages with insufficient stall marginand durability. Stalls would occur often duringtesting, causing blades and vanes to clash andclang as a result of insufficient flexural rigidity.This experience promoted the use of lower aspectratio airfoils for both stall margin and durability.

A PFRT redesign was initiated using 11 long

chord stages in titanium and A286 at a pressureratio of 9 in the same axial length. The J93 enginehad variable inlet guide vanes (IGVs), 3 front and5 rear cantilevered variable stators, an annularcombustor, convection cooled turbine bladesoperating at 2000

oF RIT and a large converging

diverging exhaust nozzle. The rear variable statorswere opened above Mach 2 to decrease overallpressure ratio and increase airflow at the higher


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Mach numbers. Extensive development of the J93resulted in producing significant turbomachinerytechnologies for all future GE engines.

Figure 32 GE J93 Mach 3 Turbojet Engine(Courtesy of GE)

In the same time period P&W was also funded todevelop high Mach engine technology. This effortled to the development of the J58 Mach 3+turbojet (Figure 33) for the twin engine Lockheed

SR-71 high altitude reconnaissance aircraft. TheJ58 engine used variable inlet guide vanes(IGVs), a 9 stage compressor, a mid stage bleedbypass, a cannular combustor, the first P&Wconvection cooled turbine rotor at 1780

oF RIT and

an afterburner. Compressor bypass bleed doorsafter stage 4 opened at high Mach decreasing thecompressor pressure ratio and dischargetemperature while increasing airflow. Thisproduced a similar result as the J93 with rearvariable stators introducing a variable pressureratio cycle. The exhaust nozzle consisted of aP&W primary section and a Lockheed blow in doorejector section for the secondary which wasintegrated into the aircraft structure. A newgeneration of advanced superalloys andprocessing was also developed to meet theextended high temperature operation. The J58was the only operational Mach 3+ engine inservice for over 30 years.

Figure 33 PW J58 Mach 3+ Turbojet EngineShown With Primary Nozzle

(Courtesy of P&W)

Core Engine Evolution

The introduction of either dual spool turbojets orturbofans redefines the traditional core engineused to produce propulsive energy (Figure 34).

The core engine now includes the HPC, LPC andfan hub (in blue) with the drive turbines HPT andLPT (in red). The fan bypass (Wf) using LPT driveturbine energy (in yellow) provides a lowerexhaust velocity for improved performance.

Figure 34 Core Engine with Dual Spools and Fan

Gas turbine evolution is best represented bycomparing the power of the core engine with theturbine rotor inlet temperature (Figure 35 circa1990). This comparison is also a guide to future

development opportunities for increasing coreengine performance. The extended core enginepower is normalized by dividing by the mass flowto eliminate the size effect representing thespecific core engine power. By 1990, the coreengine power increased steadily with turbinetemperature by more than 5 times over the earlyengines. The ideal Brayton cycle performance(100% efficiency and no cooling air) is representedby the formula showing that specific power is afunction only of turbine rotor temperature.

PW MILITARY

GE90

Figure 35 Core Engine Performance Evolution

The engines shown with losses and cooling air,produce about 30-35% less power than the ideal.It takes higher efficiency, higher temperaturematerials and improved cooling effectiveness tomove closer to the ideal performance.

The ideal specific power increases with turbinetemperature until reaching the fuel stoichiometrictemperature. The stoichiometric temperature ofhydrocarbon fuel is a function of the compressor


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discharge temperature (T3) (Figure 36). With theaverage compressor discharge temperature T3limited to 1200

oF, the stoichiometric temperature

is 4300oF and a barrier for gas turbine engines.

Figure 36 Fuel Stoichiometric Temperature vs. T3

Control Systems

The control system is the brain of the enginetranslating the throttle signal request into single ormultiple specific commands to the various engine

components. During the past 40 years, controlsystem capability has seen a remarkable evolutionfrom hydromechanical analog to multi-function fullauthority digital electronic controls (Figure 37).

Figure 37 Engine Control System Evolution

In the 60s, the hydromechanical control served asan analog computer with 5 functions involvingcombustor, afterburner fuel flow, inlet guide vane,stator vane position, bleed / cooling valve settings,exhaust nozzle position and rotor speed feedback.By the 70s, the complexity of these controls madethem both difficult and costly to maintain. Thedevelopment of a reliable digital integrated circuitat P&W allowed engineers to design electronic

trimmers for the B727 aircraft to reducehydromechanical complexity. In the mid 80s a fullauthority, single channel Digital Engine ElectronicControl (DEEC) with partial redundancy, wasdeveloped using a small hydromechanical controlas a backup for get home capability. The DEECcould self trim the engine to maintain thrust levelin flight and also made it possible to developvectoring nozzles for fighter aircraft. The mid 80sF15 demonstrator aircraft (Figure 38)with pitch

vectoring nozzles and electronic controls led anew generation of engines providing greaterfighter aircraft maneuverability and capability.

Figure 38 NASA F15 Flight Test with ElectronicControls and Pitch Vectoring Nozzles

In the 90s, the P&W designers developed a FullAuthority Digital Electronic Control (FADEC) withsufficient redundancy and reliability to eliminate

the hydromechanical backup unit. FADEC controlsnow have the ability to detect engine failures andboth isolate and accommodate them for safety,get home capability and maintenance. Dualchannel FADEC controls were qualified on the PWF119 engine (Figure 39)incorporating pitchvectoring nozzles for the F22 stealth fighter.

Figure 39 F119-PW-100 Fighter Engine with DualFull Authority Digital Electronic Controls

(Courtesy of P&W)

The F119 incorporates all key technologiesdeveloped over the past 50 years. The dualFADEC control units are mounted aft on bothundersides of the lower fan duct. All accessoriesincluding pumps, valves, oil tank and gearbox aremounted on the underside with the ability for quickmaintenance with hand tools. The low radius ratiofan blades are hollow diffusion bonded titanium

and linear friction welded to the disk making a onepiece rotor stage. The compressor blades aremachined integral with the disks to eliminatedovetail attachments reducing leakage and weight.The rotor spools are counter rotating with the highpressure rotor piggy back mounted on rollerbearings inside a flange disk on the low pressureshaft. Placing the high speed rotor roller bearinginside a race mounted on the low speed shaftreduces the normal radial clearance and provides


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Figure 43 Advanced Ducted Propulsor Engine(Courtesy of P&W)

The higher bypass ratio fan was driven by a highspeed low pressure turbine through a gear drivewith a 3.7 reduction ratio. A gear driven turbofanengine is more complex than the standard directdrive configuration but offers the potential for

achieving higher propulsive efficiency (Figure 44).

Figure 44 Gear Driven Turbofan Engine

A gear drive requires at least 98% efficiency butallows the low pressure turbomachinery to operateat higher rotational speeds reducing the number ofstages. The gear driven turbofan configurationoffers the possibility to increase overall efficiencywithout an increase in core thermal efficiency.However, since these very high bypass ratioengines produce less thrust per pound of airflow,they require a larger diameter to produce thesame thrust as a direct drive turbofan.

Summary

During the past 50 years, it took many thousandsof engineers and considerable resources to bringthe jet engine to its current state of evolution.Overall engine efficiency (thermal x propulsive) fortransports has increased from 20% for turbojets to36% for high bypass turbofans. Lessons learnedhave repeatedly shown that technology shouldlead the commitment to meet program

milestones and achieve the lowest developmentcost. However, many technologies weredeveloped out of necessity in the heat of battleduring major development initiatives suggestingmotivation, ability and perseverance are also keyingredients for success.

The Advanced Turbine Engine Gas Generator(ATEGG) initiative began in the early 60s and wasincluded in the Integrated High PerformanceTurbine Engine Technology (IHPTET) program inthe 80s. These government programs provided abenchmark in leadership for engine technologydevelopment during the past 40 years accountingfor most of the technology evolution presented.

In the case of the supersonic transport, issues ofeconomics, NOxemissions and airport noise havenot been resolved. A national initiative would berequired to overcome these remaining technicalbarriers to realize supersonic travel which was

Frank Whittles dream when he proclaimed thefuture will be more exciting than the past.

Technologies presented have been transitioned tomarine propulsion, industrial engines for powergeneration, rocket engine turbomachinery andramjets. Looking forward, candidate technologiesthat designers need for the future include:

higher temperature, non oxidizing superalloys,

very low NOx combustors for aero engines,

high temperature non-burning titanium alloys,

ductile composites with increased strain range

and

a national initiative to improve the performance

for both subsonic and supersonic aero engines.

Ultimately, such technologies need transition intoproducts to advance the state of art whileproviding superiority in defense and a positiveaeronautics balance of trade for the United States.Design is the creative artform of analysis and thesource of competitive advantage.

References:1

Midland Air Museum, Coventry Airport,Baginton, Warwickshire, U.K.

2Deutsches Museum, Bonn, Meisterwerke,Germany

3University of Southampton, HighfieldSouthampton, U.K.

4U.S Air Force Museum, Dayton, Ohio

5IbidSpanning the Globe with Jet Propulsion,

B.L. Koff AIAA Paper # 2987, April 30, 1991The History of Aircraft Gas Turbine Development

In the U.S., James St. Peter (1999)

koff gas turbine technology evolution a designers perspective

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