NASA Technical Memorandum 89368 '
`
`Analysis of an Advanced Technology
`Subsonic Turbofan Incorporating
`Revolutionary Materials
`
`Gerald Knip, Jr.
`Lewis Research Center
`
`Cleveland, Ohio
`
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`N87-22680
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`May 198?
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`ANALYSIS OF AN ADVANCED TECHNOLOGY SUBSONIC TURBOFAN INCORPORATING
`
`REVOLUTIONARY MATERIALS
`
`Gerald Knip, Jr.
`National Aeronautics and Space Administration
`Lewis Research Center
`Cleveland, Ohio 44135
`
`SUHHARY
`
`implementation of revolutionary composite materials in an
`Successful
`advanced turbofan offers the possibility of further improvements in engine
`performance and thrust-to~weight ratio relative to current metallic materials.
`The present analysis determines the approximate engine cycle and configuration
`for an early Zlst century subsonic turbofan incorporating all composite mate-
`rials.
`The advanced engine is evaluated relative to a current technology base-
`line engine in terms of its potential fuel savings for an intercontinental
`quadjet having a design range of 5500 nmi and a payload of 500 passengers.
`
`E—3542
`
`two~spool, advanced engine has an
`The resultant near optimum, uncooled,
`overall pressure ratio of 87, a bypass ratio of 18, a geared fan, and a turbine
`rotor-inlet temperature of 3085 °R. Relative to the baseline,
`the advanced
`engine yields a 22 percent
`improvement
`in cruise TSFC and a 36-percent reduct-
`ion in engine weight. Together these improvements result in a 33—percent fuel
`saving for the specified mission.
`
`For
`Various advanced composite materials are used throughout the engine.
`example, advanced polymer composite materials are used for the fan and the low
`pressure compressor (LPC).
`A Ti metal matrix composite is used for the high
`pressure compressor (HPC)
`to accommodate the higher operating temperatures.
`Ceramic composites are used for the combustor and both turbines.
`
`The advanced engine's performance includes aggressive component efficien-
`cies based on these new materials and structural changes such as swept fan and
`compressor blades. uncooled turbines,
`reduced hub tip ratios, higher blade
`loadings,
`reduced clearances, and three-dimensional design concepts.
`
`INTRODUCTION
`
`the past 40 yr, advanced turbine engines have been the pacing item
`Over
`in terms of
`the U.S. competitive edge in commercial aviation. During this
`period the specific fuel consumption (TSFC) of subsonic turbine engines has
`been reduced by about 40 percent.
`This reduction has been achieved through
`improvements in component aerodynamics, materials, and turbine cooling effec-
`tiveness.
`Improvements in materials and turbine cooling have resulted in the
`maximum turbine temperature being increased from T000 to 2600 °F. Engine
`overall pressure ratios have increased from 5 to over 38 (refs.
`1
`to 3), and
`bypass ratios from 0 (turbojet) to 7 (turbofan). Advanced metallic materials
`have also allowed tip speeds and blade loading to be increased resulting in
`fewer but more efficient stages and lighter weight components. Composite
`materials are just beginning to be used in turbine engines and then only for
`nonrotating components,
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`implementation of revolutionary composite materials (e.g.,
`Successful
`polymer, metal matrix, and high temperature nonmetallic composites)
`in an
`advanced turbofan offers the possibility of still further improvements in
`engine performance and thrust-to—weight ratios relative to current conven—
`tional materials (i.e., titanium, steel, and superalloys). Advanced composite
`materials with advanced structures will allow higher tip speeds and thinner
`blades resulting in fewer and more efficient stages. Advanced nonmetallic
`composites will allow turbines to operate uncooled at higher turbine inlet
`temperatures resulting in higher overall pressure ratios and bypass ratios,
`thereby improving performance.
`Lower material densities and,
`therefore,
`reduced blade weights will result in lower stresses and reduced engine weight.
`Advanced structures such as drum construction rather than disks will also
`
`result in lower engine weights.
`
`The purpose of this study was to (1) determine the approximate cycle and
`configuration for a turbofan engine incorporating revolutionary all-composite
`materials, and (2) evaluate the potential
`fuel saving relative to an engine
`using current
`technology (current material) for a commercial subsonic trans-
`port mission. This was done by conducting both engine cycle and flowpath
`studies.
`
`ANALYSIS
`
`Mission
`
`this study, an intercontinental quadjet having a design range of
`For
`5500 nmi and a payload of 500 passengers was assumed.
`Engines having a thrust
`of about
`lo 000 lb at Mach 0.8 and 35 000 ft would be required. This size
`engine was, therefore, considered in the present study. Sensitivity factors
`for engine performance (TSFC) and weight were used to determine changes in
`fuel consumption.
`
`Baseline Engine
`
`The baseline engine used for the study is similar to the Maximum Effi-
`ciency Energy Efficient Engine of reference 4.
`It is a two—spool engine with
`the fan and the low pressure compressor directly driven by the low pressure
`turbine.
`The engine is based on current technology. Compressor pressure
`ratios are listed in table I along with the turbine rotor~inlet temperature.
`Air for cooling the turbine is extracted at the exit of the high pressure com-
`pressor. Turbine cooling requirements for the baseline engine are based on
`the method outlined in reference 5. Compressor exit bleed requirements for
`turbine cooling are based on values for cooling effectiveness assuming advanced
`convection cooling with trailing edge ejection. Current allowable bulk metal
`temperatures were used for the vanes (2200 °R) and blades (2100 °R). Based on
`the turbine stage cooling requirements.
`the stage efficiency was corrected
`accordingly.
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`Advanced Lngines
`
`the advanced engine, a cycle study was conducted to
`— For
`Qx£l§mé£gg1.
`define an engine cycle based on thrust specific fuel consumption (TSFC) as the
`figure of merit.
`TSFC is influenced by engine cycle parameters such as over-
`all pressure ratio (OPR), bypass ratio (BPR),
`turbine rotor—inlet temperature
`(T41), component efficiencies, and component configurations.
`An in-house
`design-point study program (FACE) was used to determine design point TSFC's
`and to screen various separate flow turbofan configurations.
`Inputs for this
`program include fan and compressor pressure ratios, number of stages,
`type of
`compressor (axial or centrifugal),
`turbine rotor-inlet temperature. number of
`spools, number of turbine stages, and type of turbine cooling configuration.
`
`two levels of compressor and turbine efficiencies
`the present study,
`For
`were considered.
`one level represents current
`technology as opposed to the
`aggressive second level, which represents advanced technology. Based on dis-
`cussions with NASA component personnel and a study conducted under NASA con-
`tract (ref. 4), higher component efficiencies due to the use of advanced
`composite materials were postulated based on components having thinner blades.
`higher tip velocities, uncooled turbines,
`improved clearance control, and
`reduced hub-tip ratios in addition to making more efficient use of advanced
`three-dimensional, CFH design technology. Efficiencies for the compressors and
`turbines are determined on the basis of stage pressure ratio and work factor
`(gJ AH/N/Ufi), respectively.
`Symbols are defined in appendix A. Compressor
`and turbine efficiencies are then corrected for size effects. Turbine effi-
`ciencies are also corrected to account for clearance and turbine cooling
`effects.
`
`Advanced Components and Materials
`
`Advanced materials currently being considered to allow future turbine
`engines to operate efficiently at high temperatures and pressures are shown in
`figure l.
`
`Compressors. — Both polymer and metal matrix composites (refs. 6 to 9)
`are candidates for fan blades and the front stages of a compressor.
`For
`the
`latter stages operating at higher temperatures metal matrix and intermetallics
`are possibilities.
`The rotor may consist of a drum for retaining the blades
`(outside the scope of
`the present study).
`
`— Advanced materials are also being considered for future tur-
`combustor.
`bine engine combustors.
`These materials included oxide dispersion strengthened
`(ODS) superalloys and nonmetallic composites, such as ceramic composites and
`carbon-carbon.
`These materials would enable the combustor to be operated at
`higher temperatures with little or no cooling.
`
`improved materials and better methods of
`Turbines. - The development of
`cooling to attain higher turbine temperatures has been one of the primary meth-
`ods for achieving more efficient and higher thrust—to—weight ratio engines.
`The potential use temperatures for various materials for both vanes and blades
`are shown in figures 2 and 3.
`For both vanes and blades, nonmetallic compos-
`ites such as ceramic composites and carbon—carbon C—C (refs. 10 to l2) may per-
`mit use temperatures as high as 3460 and 4460 °R, respectively. Hith allowable
`use temperatures of this magnitude,
`there would be no need for turbine cooling.
`
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`— Advanced materials such as nonmetallic composites are also
`goggles.
`being considered for turbine engine nozzles.
`
`flowpaths for selected engines
`- Based on the cycle studies,
`Flowpath.
`were then determined using NNEP (ref.
`l3) and the NASA weight code (ref. 14).
`Thermodynamic inputs required for the weight code are determined in NNEP.
`The
`weight program determines the weight of each component
`in the engine such as
`compressors, burner,
`turbines, frames. gearbox, and accessories. Component
`weights are determined on the basis of a preliminary design approach consider-
`ing stress levels, maximum temperature, material density, geometry, stage load-
`ing, and hub-tip ratio. Based on the advanced materials of figure 1 selected
`for each component and the data presented in reference 4,
`the weight program
`input parameters for the baseline engine were updated to account for differ-
`ences between the baseline turbofan and the advanced turbofan using advanced
`composite materials.
`The material for each component was selected on the basis
`of maximum component
`temperature. Relative to a current aluminum or magnesium
`gearbox housing, a metal matrix composite may result in a stiffer housing.
`However, for this study no weight advantage was considered.
`
`- The uninstalled performances for the baseline and
`Installed performance.
`the advanced engines were corrected for nacelle drag. Based on the differences
`in installed TSFC and weight between the two engines, fuel savings were deter-
`mined using mission sensitivity factors (ref.
`l5).
`These sensitivity factors
`were determined from a mission analysis of an intercontinental,
`turbofan pow-
`ered transport having a range of 5500 nmi and a payload of 500 passengers.
`
`RESULTS AND DISCUSSION
`
`To minimize TSFC for an advanced turbofan engine providing a thrust of
`T0 000 lb at Mach 0.8 and 35 000 ft, the following parameters were considered:
`turbine rotor-inlet temperature (T41).
`fan and compressor pressure ratios,
`overall pressure ratio (DPR), bypass ratio (BPR), and number of stages for a
`two-spool engine.
`
`Effect of Advanced Engine Design Parameters on Performance
`
`turbine inlet
`- The level of
`Turbine rotor-inlet temperature (T41).
`temperature can affect not only engine performance (TSFC), but also engine
`size and,
`therefore weight.
`The effect of T4] on TSFC for a range of 0PR's
`is shown in figure 4.
`For
`these advanced engines,
`the turbines are uncooled.
`The pressure ratio for the fan (l.55(bypass)/l.4(core)) and the low pressure
`compressor (LPC) were held constant while the pressure ratio of the high pres-
`sure compressor
`(HPC) was varied to achieve the specified OPR.
`The number of
`axial stages for the LPC and the HPC were fixed at 3 and ll, respectively.
`Although fewer stages for the HPC could be used for the lower pressure ratios,
`this would result
`in lower compressor efficiencies and,
`therefore, a higher
`TSFC.
`Two axial stage turbines were used for the high pressure turbine (HPT)
`and five for the low pressure turbine (LPT). Bypass ratio was optimized for
`each cycle with respect
`to TSFC. Turbine inlet temperatures (T41) between
`2760 and 3085 °R have a small effect on TSFC (fig. 4) due in part to the
`aggressive component efficiencies for the advanced engine. However,
`increas-
`ing the OPR from 40 (current
`technology)
`to 100 results in about an 8.5 percent
`decrease in TSFC.
`The effect of turbine temperature on specific thrust (based
`
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`from 2760 to 3085 °R
`Increasing T4]
`is shown in figure 5.
`on core flow)
`A higher T41 results
`results in a 35 percent
`increase in specific thrust.
`in a somewhat higher specific thrust but also a higher TSFC.
`Since fuel con-
`sumption is more sensitive to TSFC than engine size, and therefore weight, a
`temperature of 3085 °R was selected.
`
`— The effect of fan pressure ratio (and therefore
`[an pressure ratio.
`BPR) on TSFC is shown in figure 6. Again,
`the LPC pressure ratio was fixed at
`2 and the HPC pressure ratio varied. Decreasing the fan pressure ratio from
`1.78 to 1.33 results in about a 7 percent decrease in TSFC.
`Lower fan pressure
`ratios with their associated lower tip speeds can be accommodated by means of
`a gearbox without penalizing turbine efficiency or turbine weight.
`For an OPR
`of 100, figure 7 shows the effect of BPR on TSFC.
`The optimum BPR increases
`from 18 to 28 as fan pressure ratio is decreased from 1.55 to 1.33. Corre-
`spondingly, TSFC decreases by 2.5 percent. However, since the higher BPR
`would result in a significant increase in engine size and.
`therefore drag, a
`fan pressure ratio of 1.55 was selected for the advanced engine.
`
`Low pressure compressor. - whereas the fan is connected to the low pres-
`sure turbine by means of a gearbox.
`the LPC is directly connected to the tur-
`bine. Operating the LPC at the same speed as the turbine increases the tip
`speed, resulting in fewer stages for the same pressure ratio.
`Increasing the
`pressure ratio of the three—stage LPC from 2 to 2.8 (fig. 8) has almost no
`effect on TSFC. Although the pressure ratio of the LPC does not affect per-
`formance, it may have to be modified to change the required pressure ratio of
`the HPC when one considers component diameters for the flowpath.
`To reduce
`the pressure ratio of the HPC, an initial pressure ratio of 2.8 for the LPC
`was selected for the advanced engine. Based on the relatively small penalty
`associated with reducing the OPR from 100 to 87 (fig. 6), an OPR of 87 was
`selected for the advanced engine.
`
`technology all—axial HPC's have
`- Current
`High Pressure Compressor (HPC).
`an average stage pressure ratio of about 1.3. Using this as the average stage
`pressure ratio,
`the HPC for an OPR of 60 and 87 would require ~11 and 13
`stages, respectively.
`For an OPR of 81,
`the effect on TSFC of reducing the
`number of HPC stages from 13 to 10 is minor (fig. 9). However, 10 or 11 stages
`may be advantageous in terms of engine length, weight, and cost.
`As a result
`the average stage pressure ratio would increase to 1.39 and 1.35, respectively.
`These increased pressure ratios can be achieved by means of higher tip speeds
`and/or increased blade loading.
`
`In addition to the all—axial high pressure com-
`—
`Axial—centrifugal HPC.
`pressor previously discussed, an axial—centrifugal HPC was also considered
`(fig. 10).
`figure 10 shows the optimum pressure ratio split in terms of TSFC
`for various axial-centrifugal compressors having an overall pressure ratio of
`87. At
`these high pressure ratios, it may be necessary to replace some of the
`latter axial stages (because of minimum blade height limitations of ~l/2 in.)
`with a centrifugal stage.
`TSFC decreases as the number of axial stages is
`increased from 3 to 5 due to the higher efficiency of the axial stages.
`The
`optimum pressure ratio for the centrifugal stage decreases from about 6 to 4
`as the number of axial stages increases from 4 to 5.
`The average pressure
`ratio for the axial stage is ~1.47. Because of the slightly lower TSFC
`(fig. 9) for the all—axial compressor, it was selected for the advanced engine.
`
`— The actual stage mean velocity of the HPT
`High pressure turbine (HPT).
`depends on the RPM of the HPC and/or the turbine work factor (gJ AH/N/Ufi).
`5
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`shows
`Figure 11
`the order of 1400 ft/sec.
`Currently mean velocities are of
`the effect on TSFC of increasing the mean velocity from 1500 to 1600 ft/sec
`for a HPT consisting of 2 to 4 stages. Based on these results the choice as
`to mean velocity will depend more on the engine flowpath and weight than TSFC.
`
`the mean velocity of
`— For direct drive fans,
`Low pressure turbine (LPT).
`current LPT's is about 900 ft/sec. with a geared fan, such as would be used
`on the advanced engine,
`the mean velocity of the LPT could be increased,
`thereby,
`reducing the stage work factor and increasing the turbine efficiency.
`Figure 12 shows the effect on TSFC of
`increasing the mean velocity and varying
`the number of stages.
`The required expansion ratio for the cycle is about
`20.
`Thus five to possibly as many as seven turbine stages may be required.
`Again the flow path must be considered before selecting the number of LPT
`stages and/or the mean velocity.
`
`Advanced Engine Preliminary Definition
`
`Based on the initial cycle and configuration studies the two-spool, sepa-
`rate flow, advanced turbofan at the cruise design point is defined as follows:
`
`1.55/1.4
`
`Fan pressure ratio:
`Bypass ratio:
`18
`P/P = 2.8
`LPC three-axia1 stages:
`P/P = 22.2
`HPC e1even—axia1 stages:
`Turbine rotor-inlet temperature:
`T4]
`HPT:
`three-axial stages
`LPT:
`five-axial stages
`Bleed: 0.0
`
`= 3085 °R
`
`The number of stages for the various components was considered again in the
`flowpath portion of the study.
`
`Advanced Engine Flowpath
`
`- To this point. cycle studies have
`Qygje and geometrical_tradeoffs.
`identified the pressure ratios,
`turbine temperature. bypass ratio. and the
`configuration resulting in about
`the minimum TSFC for the advanced engine.
`However,
`to optimize the engine in terms of TSFC and weight,
`the engine flow-
`path must be considered to identify possible tradeoffs required of the cycle
`to achieve an acceptable gas path geometry.
`For example,
`to reduce the number
`of stages of compression, blade loading. tip speed, and stress levels must be
`considered along with pressure ratio and efficiency.
`The compressor in turn
`affects the dimensions of
`the turbine based on the loading parameter resulting
`from the cycle study.
`A short study of
`the flowpath involving various cycle
`and geometrical
`tradeoffs was,
`therefore, undertaken. These tradeoffs are
`discussed in detail
`in appendix 8 and summarized below.
`
`- Based on the flowpath
`Compar1;9Q_of advanced and baseline engines.
`studies, various component configurations, tip speeds, and loadings resulting
`from the cycle study were modified.
`The number of stages for the LPC was
`decreased from three to two in order to save weight.
`The number of HPT stages
`was
`reduced from three to two while the tip velocity was increased by about
`200 ft/sec.
`As a result the loading parameter decreased from 2.0 to 1.54.
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`In addition the loading parameter for the LPT was increased from 1.52 to 1.86.
`These changes reduced the length of the transition duct between the HPT and
`the LPT in addition to reducing engine weight.
`The resulting flowpath for the
`advanced turbofan engine is shown in figure 13.
`For purposes of comparison,
`the flowpath for the baseline engine is shown in figure 14.
`The fan diameter
`is smaller for the baseline because of its lower bypass ratio. Core radii for
`the advanced engine are,
`in general, smaller due to the higher tip velocities.
`Engines having higher bypass ratios and a direct drive fan usually have a
`longer transition section between the turbines so as to achieve higher effi-
`ciencies for the required work loads. This can be seen by comparing figures
`13 and 14.
`
`Parameters resulting from the cycle and the flowpath studies for the
`advanced engine are compared in tables II and III with those for the baseline
`engine. Table II compares various geometric parameters,
`the number of stages,
`and the number of rotor blades. Table III compares the cycles, and flowpath
`temperatures and pressures. Engine OPR and turbine rotor—inlet temperature
`(141)
`increased from 42.5 and 2175 “R for the baseline engine to 87 and
`3085 °R for the advanced turbofan. Although the advanced engine operates at a
`higher T41. it is uncooled due to the higher use temperature of the advanced
`materials.
`The higher OPR is achieved in one less stage of compression.
`Because of
`the higher operating pressure ratios,
`temperatures exiting the
`various components are in general higher.
`
`Based on temperatures associated with the LPT of the advanced engine. a
`current material (density of 0.3 lb/in ) could be considered for this compo-
`nent. However, an advanced material such as a ceramic composite (density of
`0.115 lb/in.3) would reduce the weight of
`the engine by 7.5 percent.
`figure 15. Advanced materials considered for the other components of the
`advanced engine are listed in table IV.
`
`Figure 16 compares the uninstalled performance of the selected advanced
`turbofan with that of
`the baseline engine.
`On an uninstalled basis the
`advanced engine results in a 23 percent reduction (improvement)
`in TSFC.
`
`The uninstalled performance for both engines was then corrected for
`nacelle drag based on the advanced nacelle geometry (ref. 4) specified in
`table 5.
`The advanced nacelle is compared with a current turbofan nacelle in
`figure 17.
`The pressure drag was considered to be equal to 50 percent of the
`friction drag.
`On an installed performance basis,
`the TSFC of the advanced
`turbofan engine is 22 percent
`lower than the baseline.
`
`lighter than the
`the advanced engine is 36 percent
`In terms of weight,
`baseline (fig. 18). This is due to both the lower densities of the advanced
`materials and the increase in the turbine rotor-inlet temperature.
`The higher
`turbine temperature decreases the size of the engine for a given thrust and,
`therefore,
`its weight.
`The advanced engine with its improved TSFC and reduced
`weight results in a 33.5 percent fuel savings relative to the baseline for an
`intercontinental quadjet having a range of 5500 nmi and a payload of 500
`passengers (fig. 19).
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`CONCLUSIONS
`
`The selected cycle resulting from using advanced materials for an advanced
`technology (2010), subsonic turbofan engine has an overall pressure ratio of
`87. a bypass ratio of 18, a gear driven fan, and a turbine rotor—inlet
`tempera-
`ture at cruise of 3085 °R.
`Polymer and metal matrix composites are used in the
`cool sections of
`the engine and ceramic composites in the hot sections.
`As a
`result the turbines are uncooled.
`
`In addition to the higher turbomachinery component efficiencies postulated
`from the use of advanced composite materials,
`these materials are estimated
`(based on a clean sheet design approach)
`to result in the following improve-
`ments for an advanced engine with a geared fan relative to a current technology
`baseline engine.
`
`improvement
`1. A 23 percent
`condition (H.8/35K ft)
`2. A 36 percent decrease in engine weight
`3. A 33.5 percent fuel savings for an intercontinental
`range of 5500 nmi, and a payload of 500 passengers
`
`in uninstalled engine TSFC at the cruise
`
`transport having a
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`
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`
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`
`A Method to Estimate Height and Dimensions of
`Onat, E.; and Klees, G.H.:
`Large and Small Gas Turbine Engines.
`NASA CR-159481, 1979.
`
`Energy Efficient Engine Preliminary Design and Integration
`Gray, D.E.:
`Study.
`(PHA-5500-18, Pratt and Whitney Aircraft; NASA Contract NAS3—20628)
`NASA CR—13S396, 1978.
`
`9
`
`GE-1019.010
`
`UTC—2010.0l0
`
`

`
`TABLE I.
`
`- BASELINE TURBOFAN ENGINE
`H 0.8/35 000 ft
`
` Fan pressure ratio
`
`Bypass ratio
`
`
`Low pressure compressor P/P
`High pressure compressor P/P
`
`
`Combustor AP/P
`
`
`Turbine rotor—inlet temperature,
`Turbine cooling air, percent
`
`
`“R
`
`TABLE II.
`
`- COMPARISON OF FLOH PATH PARAMETERS
`
`Baseline
`
`Advanced
`
`Fan
`Hub/tip ratio
`Blade solidity
`First blade AR
`Blade material density
`Corrected tip velocity.
`Number of rotor blades
`Number of stages
`
`fps
`
`LPC
`Hub/tip ratio
`Blade solidity
`First blade AR
`rpm
`Maximum rotor speed,
`Corrected tip velocity, fps
`Blade material density
`Number of rotor blades
`Conmressor design type
`Number of stages
`
`HPC
`Hub/tip ratio
`Blade solidity
`First blade AR
`rpm
`Max rotor speed,
`Corrected tip velocity, fps
`Blade material density
`Number of rotor blades
`Compressor design type
`Number of stages
`
`HPT
`Turbine loading parameter
`gJ an/N/vfi
`Blade solidity
`First blade AR
`Blade material density
`Number of rotor blades
`Turbine design type
`Number of stages
`
`LPT
`Turbine loading parameter
`Blade solidity
`First blade AR
`Blade material density
`Number of rotor blades
`Turbine design type
`Number of stages
`
`0.49
`0.8
`1.8
`20 212
`1392
`0.13
`548
`Const. mean radius
`11
`
`1.54
`
`0.25
`2
`0.115
`84
`
`2
`
`Const. hub
`
`Const. tip
`
`.
`
`.
`
`.
`
`.
`.
`
`Const. mean
`
`2.38
`0.44
`3
`0.3
`404
`Const. hub radius
`5
`
`10
`
`GE-1019.011
`
`UTC—2010.0ll
`
`

`
`TABLE III. - ENGINE COMPARISON
`
`Cycle parameters H 0.8/35K ft
`
`Fan
`P/P
`
`Bypass
`ratio
`
`LPC
`P/P
`
`HPC
`P/P
`
`7.2
`18.1
`
`1.84
`2.8
`
`14
`22.2
`
`42.5
`87
`
`T41,
`°R
`
`2775
`3085
`
`
`
`
`
`Baseline
`Advanced
`
`Flow path temperatures and pressures
`
`Inlet Bypass
`
`HPC
`LPC
`exit exit
`
`HPT
`HPT
`inlet exit
`
`LPT
`exit
`
`633
`15.9
`
`1264
`
`8.8
`
`676
`20.6
`
`1109
`
`4.8
`
`
`
`
`
`
`
`
`
`
`Baseline
`TI,
`°R
`PT, psi
`Advanced
`T1.
`°R
`PT. psi
`
`445
`5.3
`
`445
`5.3
`
`522
`8.7
`
`507
`8.1
`
`TABLE IV. — ADVANCED ENGINE MATERIALS
`
`Component
`
`Material
`
`Density,
`lb/in3
`
`T1 metal matrix composite
`
`Polymer composite
`Polymer composite
`Ti metal matrix composite
`Ceramic composite
`Ceramic composite
`Ceramic composite
`
`TABLE V.
`
`— NACELLE GEOHETRY
`
`
`
`
`(L/D)max
`Fan cowl
`thickness ratio,
`Fan cowl
`9max/°tip- 0
`Fan hub/fan tip, Dhub/Dtgp
`Fan specific flow, K,
`lb/sec/ft?
`Diameter ratio of scrubbed surface
`Fineness ratio of scrubbed surface
`
`
`
`
`11
`
`GE-1019.012
`
`UTC-2010.012
`
`

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`FIGURE 3.
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`
`
`
`\
`
`12
`
`GE-1019.013
`
`UTC—20l0.0l3
`
`

`
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`FIGURE 8.
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`
`18
`
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`
`GE-1019.014
`
`UTC—20l0.0l4
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`FIGURE 7. - OPTINUH BYPASS RATIO IN TERNS W ENGIIE
`PERFORMANCE Fm RANII U’ FAN PRESSURE RATIOS.
`
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