Advanced Core Technology: Key to Subsonic Propulsion Benefits

A study was conducted to identify the potential performance benefits and key technology drivers associated with advanced cores for subsonic high-bypass turbo-fan engines. Investigated first were the individual sensitivities of varying compressor efficiency, pressure ratio and bleed (for turbine cooling); combustor pressure recovery; and turbine efficiency and inlet temperature on thermal efficiency and core specific power output. Then, engine cycle and mission performance benefits were determined for systems incorporating all potentially achievable technology advancements. The individual thermodynamic sensitivities are shown over a range of turbine temperatures (at cruise) from 2900 to 3500 °R and for both constant (current technology) and optimum (maximum thermal efficiency) overall pressure ratios. It is seen that no single parameter by itself will provide a large increase in core thermal efficiency, which is the thermodynamic parameter of most concern for transport propulsion. However, when all potentially achievable advancements are considered, there occurs a synergism that produces significant cycle and mission performance benefits. The nature of these benefits are presented and the technology challenges associated with achieving them are discussed.

Core thermal efficiency is herein defined as the ideal power available from the engine core (i.e., the output from the idea] power turbine) as a fraction of fuel heating power.
The pressure ratio across the power turbine was adjusted so the core stream would produce no net thrust.
The supercharger of the cycle model represents rne compression provided by the inner (i.e., core) section of a fan or by a low-spool compressor, while the low-spool turbine provides only the power needed by the _upercharger.
Engine For the limited range of parameter variations studied, the sensitivities were all approximately linear.
Thus, average sensitivities of core thermal efficiency and specific power output were determined as tl_own in Fig. 5 for constant overall pressure ratio of 37 and i_ Fig. 6 for optimum overall pressure ratio <i.e., at maximum thermal efficiency).
The six independent parameters for which the sensitivities were determined are shown at the bottoms of ti)e bars.
At the tops of the bars are the units for oxpressing the sensitivities.
The plus or minus sign inside the bar represents the direction of the sensitivity slope.
The gray area, where it exists, at the top <>F a bar shows the effect of turbine inlet temperature ,_n the sensitivity.
In all cases, except for the effect <>f OPR on thermal efficiency at constant pressure ratio, where the reverse is true, the top of the gray area is for 2900 °R while the bottom of the gray area is for 3500 °R.
No gray area at the top of a bar indicates the absence of a temperature dependency.
From Figs. 5 and 6 we see that thermal efficiency and/or specific power benefits are available to some _,xtent from all the cycle and component parameters studied.Burner pressure recovery does not offer much potentia] benefit because of the relatively small sensitivities and the high values, better than 95 percent clready being achieved.
The sensitivities to turbine and compressor efficiencies are higher, about twice as high, at optimum pressure ratio than at the constant aressure ratio.This occurs because output power is becoming a smaller difference between two larger work terms as pressure ratio increases.
Turbine inlet temperature affects specific power conslderably more than it does thermal efficiency.
So does compressor bleed since increasing bleed is effectively the same as reducing turbine inlet temperature.
Both thermal efficiency and specific power display maximum values as pressure ratio varies, the pressure ratio for maximum thermal efficiency being much higher than that for maximum specific power.
As seen from the sensitivity slope directions indicated in Fig. 5, the constant overall pressure ratio of 37 is greater than that for maximum specific power but less (also see Fig. 4) than for maximum thermal efficiency.
Thus, increasing pressure ratio from current values offers increased thermal efficiency but at the expense of reduced specific power.
Since thermal efficiency is of primary importance for subsonic transport propulsion systems, increased pressure ratio should be explored.However, it should be noted that there is no single parameter that by itself will provide a large increase in thermal efficiency.With improved efficiencies for the core compressor and turbine, the thermal efficiency improvement increases to 14 to 18 percent at a pressure ratio of about lO0.The cross-hatched band in Fig. 7 represents varying degrees of optimism concerning component efficiency improvements.

ENGINE CYCLE AND MISSION PERFORMANCE
These component efficiency gains would come not only from aerodynamics improvements, but also from the higher rotative speeds and reduced Teakages enabled by technology advancements in materials, structures, bearings, and seals.
Finally, if uncoo]ed turbine temperature can increase to 3460 °R, further small improvements in thermal efficiency are thermodynamically achievable, but require very high pressure ratios, values beyond ]00.
A major point to be drawn from Fig. 7 is that neither pressure ratio by itself, advanced component technology by itself nor turbine temperature by itself provides a large increase in thermal efficiency.
However, increasing the engine overall pressure ratio enables significant improvements to be made through synergistic coupling with increased turbomachine efficienties, reduced cooling bleed, and increased turbine temperature.
Through this synergism, the thermal efficiency improvement is at least double that which would be determined from a linear combination of individual improvements.
Asthe sensitivity analysisshowed, corespecific power outputis stronglydependent on turbinetemperature.Figure8, whichpresents the changes in corespecific power with advanced coretechnology, displaysa band to show the specificpower requiredfor ultra-highbypass turbofans capable of providinglarge improvements in propulsive efficiencywithoutexcessive engine size andweight.Asseen, turbinetemperatures on theorder of 3460 °Rare needed to meet thesepower requirements because the coreflow will be5 percent or lessof the total air flow.
Thepotential benefitsof advanced low-spool and nacelletechnologies are shown in Fig. 9. Thetop curveis for the idealizedcaseof nonacellelossand reflects the increasing propulsive efficiencywith decreasing fan pressure ratio.Nith currenttechnology nacellelossesincluded, oneobtainsthe bottom curve, whichshows that the nacellelossesmore thanoffset the increases in propulsive efficiencyas bypass ratio andengine diameter increase.Foradvanced nacelle technology (i.e., shortthin streamlined contour), performance is improved to the cross-hatched band.Finally, the dashed curverepresents the inclusionof a 3 percent gain in fan efficiency(with respect to the topof the band).Onthe basisof this analysisof the impact of low-spool andnacelletechnology advancements, it appears that an8 to I0 percent reductionin SFC is achievable.