Christopher E. Hughes, Dale E. Van Zante, and James D. Heidmann

Glenn Research Center, Cleveland, Ohio

Aircraft Engine Technology for 

Green Aviation to Reduce Fuel Burn 

NASA/TM—2013-217690

November 2013

AIAA–2011–3531
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Christopher E. Hughes, Dale E. Van Zante, and James D. Heidmann

Glenn Research Center, Cleveland, Ohio

Aircraft Engine Technology for 

Green Aviation to Reduce Fuel Burn 

NASA/TM—2013-217690

November 2013

AIAA–2011–3531

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2013-217690 1 

Aircraft Engine Technology for 

Green Aviation to Reduce Fuel Burn 

Christopher E. Hughes, Dale E. Van Zante, and James D. Heidmann

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

The NASA Fundamental Aeronautics Program Subsonic Fixed Wing Project and Integrated Systems 

Research Program Environmentally Responsible Aviation Project in the Aeronautics Research Mission 

Directorate are conducting research on advanced aircraft technology to address the environmental goals 

of reducing fuel burn, noise and NOx emissions for aircraft in 2020 and beyond. Both Projects, in 

collaborative partnerships with U.S. Industry, Academia, and other Government Agencies, have made 

significant progress toward reaching the N+2 (2020) and N+3 (beyond 2025) installed fuel burn goals by 

fundamental aircraft engine technology development, subscale component experimental investigations, 

full scale integrated systems validation testing, and development validation of state of the art computation 

design and analysis codes. Specific areas of propulsion technology research are discussed and progress to 

date.

Introduction

In 2006, an Executive Order was signed by the President of the United States issuing a National 

Aeronautics Research and Development Policy and Plan (Ref. 1), part of which was aimed at addressing 

environmental concerns in the U.S. Aviation System. Goals and objectives were established for 

reductions in fuel burn, noise and NOx emissions for future aircraft to address their environmental 

impact. The NASA Aeronautics Research Mission Directorate (ARMD) established the Fundamental 

Aeronautics Program (FAP) to develop advanced technologies to address these goals. Four distinct 

projects were created, one of which was the Subsonic Fixed Wing Project (SFW), part of whose mission 

was to develop advanced propulsion technologies for subsonic aircraft in a five to fifteen year timeframe, 

also known as N+1 to N+2 generation. Today, the SFW Project, together with research partners from 

Industry, Academia, and other Government Agencies, is focusing on developing technologies for the N+3 

generational timeframe, or beyond 2025 (Refs. 2 and 3). These are fundamental technologies in the 

Technology Readiness Level (TRL) of 1 to 3 range (definition of Technology Readiness Levels are given 

in the Appendix) that show potential benefits to address the Project environmental goals when 

incorporated to systems level applications. The SFW Project is also tasked to develop and validate

computer prediction tools at both the component and system level application for future application to 

advanced aircraft system designs.

The NASA ARMD subsequently also established the Integrated Systems Research Program (ISRP) to 

carry forward to large scale, integrated system demonstrations the technologies developed under SFW 

that show potential benefit for addressing the environmental goals. The Environmentally Responsible 

Aviation Project (ERA) was created in 2009 to develop those integrated system demonstrations to the 

TRL range of 4 to 6 in the 2020 to 2025 timeframe, or N+2 generation of aircraft systems (Ref. 4). In 

order to accomplish these complex demonstrations, ERA too relies strongly on successfully meeting the 

environmental goals by conducting research with Industry, Academia and Government Agencies through 

cost share partnerships, research collaborations, and NASA Research Announcements. In Figure 1, the 

Subsonic Transport System Level Metrics are shown, outlining the environmental fuel burn, noise and 

emissions reduction goals, the time frame for achieving them, and the baseline defined for comparing 

with new technologies. And in Figure 2, the fuel burn savings potential using advanced aircraft 
NASA/TM—2013-217690 2 

technologies available in 2020 (N+2) for several different aircraft are shown. It is predicted that advanced 

engine technology will contribute between about 15.3 and 18.5 percent of the total savings, depending on 

airframe type, compared with a large twin aisle aircraft. 

Figure 1.—NASA Subsonic Transport System Level Metrics. Fuel Burn goals for the SFW and ERA Projects are 

circled in red.

Figure 2.—Projected fuel burn savings using advanced 2020 ready aircraft technologies, and the contribution of 

each technology to total fuel savings. Advanced engine contribution to the fuel burn reduction for are circled for 

each aircraft type.

ISRP ERA Goals FAP SFW Goals

Reference Fuel Burn = 279,800 lbs

-15.9%

HWB with Composite

Centerbody

-3.2%

-5.5%

Embedded Engines with

BLI Inlets

LFC (Centerbody)

-8.7%

-3.5%

-3.6%

-15.3%

-9.9%

Composite Fuselage

Composite Wings & Tails

PRSEUS

Advanced Engines

HLFC (Wings, 

Tails, Nacelles)

Riblets, Variable TE Camber, 

Increased Aspect Ratio

Subsystem 

Improvements

-2.0% Composite Wings & Tails

-2.6% PRSEUS

-18.5% Advanced Engines

-8.6% HLFC (Outer Wings 

and Nacelles)

Riblets, Variable TE Camber

Subsystem 

Improvements

-15.9%

-1.7%

-2.3%

-16.0%

-7.5%

Riblets, Variable TE Camber

Subsystem Improvements

Technology Benefits Relative to Large Twin Aisle (Reference: 777-200LR “like” Vehicle)

Advanced "tube-and-wing” HWB300 HWB300 + more accelerated tech maturation

Fuel Burn = 159,500 lbs

-120,300 lbs (-43.0%) Fuel Burn = 140,400 lbs

-139,400 lbs (-49.8%) Fuel Burn = 128,500 lbs

-141,300 lbs (-54.1%)

-0.7%

-1.3%

-1.1%-1.1%

HWB with Composite

Centerbody

Composite Wings & Tails

PRSEUS

Advanced Engines

HLFC (Outer Wings 

and Nacelles)

-1.0%

-0.9%
NASA/TM—2013-217690 3 

This paper will provide overview of the advanced technology concepts related to engine cycle and 

power core technologies currently being investigated under the NASA SFW and ERA Projects.

Engine Design Considerations

For gas turbine based propulsion, the amount of thrust produced is the product of the mass of air per 

unit time and the change in velocity which is imparted to that air. It is known from this process that for a 

given amount of thrust it is more efficient to make a smaller change on a large volume of air versus 

having to make a large velocity change to a smaller volume of air; higher efficiency means less fuel 

burned. For these reasons, the direction of turbofan technology has been moving toward higher engine 

bypass ratios (BPR) and lower fan pressure ratios (FPR). Figure 3 illustrates the trend in fuel burn with 

BPR and FPR that as the BPR increases and the corresponding FPR decreases, the amount of fuel burned 

decreases. However, there is a downside to continuing to decrease the FPR and increase the BPR, which 

is the fan diameter must get larger to produce that same amount of thrust. A larger fan also means a larger 

engine nacelle diameter, creating additional drag at higher flight speed conditions, meaning more thrust is 

needed to overcome the higher drag. 

One other aspect affecting high bypass ratio designs is engine core size. Since noise is also a 

consideration in high bypass fan designs, fan speed must be kept as low as possible to reduce its noise 

signature. However, lower fan speeds mean lower compressor and turbine component speeds in the 

engine core as well, since a common drive shaft connects the core components and the fan. A specific 

amount of power is needed to drive the fan for a given thrust, and so as the fan design speed drops so the 

core components must get larger in size, both in number of stages and in stage diameter, to provide the 

needed power since the core component speeds cannot increase. Larger components mean more weight, 

and more fuel burn to carry that weight around. Aircraft system design studies for two different types of 

aircraft configurations are given in References 5 to 7, illustrating the impacts of different engine cycles on 

aircraft performance.

Figure 3.—Effect of fan size, bypass ratio and fan pressure ratio on fuel burn, and benefit of advanced Ultra High 

Bypass engine cycle technology.
NASA/TM—2013-217690 4 

The result of these design constraints are shown in the fuel burn trend line in red in Figure 3. It shows 

that a practical limit is finally reached as the trend line reaches a low point and then turns up again as the 

fan size increases, meaning more fuel burned. The fan size has gotten so large that the added weight of 

the fan and nacelle, the additional nacelle drag, and the increase in core weight overcomes the fuel 

advantage of the higher BPR and lower FPR design. What is needed is a paradigm shift in propulsion 

technology to extend the fuel burn advantage trend line with increasing BPR. NASA SFW and ERA 

Projects along with their research Partners in Industry, Academia and Other Government Agencies have 

been conducting advance propulsion technology research to accomplish that goal.

Engine Technologies

The following sections will describe the aircraft engine technologies being explored by the NASA 

SFW and ERA Projects to achieve the installed fuel burn metrics described in Figure 1. 

Geared Turbofan Technology

As was shown in Figure 3, a minimum point is reached in the fuel burn trend line for a given engine 

cycle where increasing the engine BPR beyond a certain point leads to negative results with higher fuel 

burn as the increase in engine weight and size overcome the benefits of a high bypass ratio engine cycle. 

With current technology, this is a result of the conventional high bypass engine cycle known as direct 

drive, where the fan and core components are rotating on the same shaft. In the direct drive engine, the 

fan and core components (compressor and turbine) operate at the same rotation speed; thus the system is 

constrained by the component with the lowest speed, which is the fan in the propulsor. The core must run 

at slower, less efficient speeds for a lower speed fan design. As discussed earlier, this means more 

compressor and turbine stages are needed to produce the power required to drive the fan and get the 

required thrust at a given operating speed, resulting in more engine weight. In Figure 3, the answer to 

getting lower fuel burn is to shift the point of maximum benefit further down on fuel burn trend line by 

introducing advanced technologies to achieve a paradigm shift. This means the fan BPR and thus 

efficiency continues increasing; but it also means the fan PR decreases while the fan diameter increases

for the required thrust. The Geared Turbofan (GTF) engine cycle introduced by Pratt & Whitney (P&W) 

is one of these advanced technologies that enable this paradigm shift. In the GTF architecture, the fan and 

the core components are separated by a gear system. The gear system allows the fan and the core to 

operate at different, more efficient speeds. The core can then operate more efficiently and produce a given 

thrust level at the fan with fewer compressor and turbine stages compared to a direct drive engine, thereby 

reducing the engine weight and the fuel required to carry that weight around on the aircraft. The 

advantages of the GTF engine cycle in terms of fan efficiency compared with the conventional turbofan 

cycle is shown in Figure 4. 

The first generation GTF was developed by P&W in partnership with NASA (Refs. 8 to 14). Scale 

model wind tunnel tests to validate predicted aerodynamic, acoustic and aeroelastic design characteristics 

were conducted in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel in 2006 Figure 5, top) 

(Ref 15). The first full scale engine demonstration of the GTF engine cycle technology was successfully 

completed by P&W in 2008 (Figure 5, bottom), and followed shortly thereafter by the first flight test 

demonstration of the GTF by P&W in the same year (Figure 5, middle). The first generation GTF is 

considered an N+1 technology (TRL 6 by 2015) which P&W plans to enter into service with aircraft 

manufacturers in 2013. In Figure 6, the technology areas investigated collaboratively by NASA and P&W 

in the developing of the first generation GTF are shown and include the low speed, low pressure ratio fan, 

fan gear system, low emissions combustor, and high speed, compact, low pressure spool. These 

technologies allowed the engine BPR to reach the Ultra High Bypass (UHB) ratio range of 12 or greater, 

while allowing the FPR to be reduced to between 1.3 and 1.4 to achieve higher fan efficiencies. In Figure 

7 (left), the projected reductions in noise, emissions and fuel burn (a fuel burn reduction of 15 percent) 

that can be achieved by the GTF engine cycle is compared with a current technology A320 aircraft and 

V2500 engine combination. 
NASA/TM—2013-217690 5 

Figure 4.—Trend in Fan Propulsive Efficiency with Fan Pressure Ratio, and range of application for 

aircraft propulsion cycles.

Figure 5.—First generation GTF technology milestones.

Better

Fuel

Burn

Worse

Fuel

Burn

First P&W flight demonstration of GTF engine in 2008 

First P&W GTF full scale 

engine demonstration in 2008

Scale model GTF fan test in NASA Glenn 

9’x15’ low speed wind tunnel in 2006
NASA/TM—2013-217690 6 

Figure 6.—NASA/P&W areas of research collaboration for the GTF engine cycle.

Figure 7.—Projected reductions in fuel burn and other environmental emissions with the P&W GTF 

engine cycle.

Once again, however, a maximum benefit point will be reached for this level of technology as the 

design BPR continues to increase and the FPR decreases. A second paradigm shift in technology is 

needed to extend the beneficial fuel burn tend line (red trend line in Figure 3) to significantly contribute 

toward achieving the ERA N+2 system level metrics environmental goals (Figure 1). NASA and P&W 

are again partnering to investigate a second generation of GTF propulsor (fan, stators, and nacelle) 

technologies to help meet those goals, enabling a BPR up to 18 and a FPR between 1.25 and 1.3 to be 

achieved. Figure 7 (right) shows the projected fuel burn benefits the second generation GTF will provide, 

which is a 25 to 30 percent reduction compared to a current V2500/A320 aircraft combination. Current 

research plans show the start of a collaborative, second generation GTF fan model test with advanced 

propulsor technologies in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel in late 2011. 

Fan Drive Gear System

Gear Ratio ~ 3

Low-Speed

Low-Pressure 

High-Bypass-Ratio

Fan

High-Speed,

Compact,

Low Spool

Low-Emissions Combustor

P&W GTF Gen 1 (2013 EIS)

NOISE -20 EPNdB

(cum margin to Ch 4)

LTO NOX -60%

(below CAEP 6)

FUEL BURN -15% 

(rel to A320/V2500)

Projected Based on

Demonstrated Technology

P&W GTF Gen 2 (2020 to 2025 EIS)

NOISE -25 EPNdB

(cum margin to Ch 4)

LTO NOX -75%

(below CAEP 6)

FUEL BURN -25% to 30%

(rel to A320/V2500)

Projected Based on

Demonstrated Technology
NASA/TM—2013-217690 7 

Open Rotor Technology 

In the mid 1980s, NASA in partnership with several U.S. Aerospace companies (Refs. 16 to 18) 

investigated a radical new propulsion technology for aircraft known as Open Rotors. As shown in 

Figure 4, this propulsion technology held the potential for large increases in efficiency, and therefore 

significant reductions in fuel burn, compared with conventional turbofans. The predicted performance of 

Open Rotor propulsion systems (Ref. 19) showed significant fuel burn benefits even then compared with 

current high bypass ratio turbofan engines (Figure 8).

As discussed earlier under Engine Design Considerations, propulsive thrust is achieved for all aircraft 

engine cycles by increasing the kinetic energy, or velocity, of a specific volume of air. The amount of air 

and the increase in velocity needed are the parameters that determine the work required to a produce a 

given amount of thrust. The relationship between the amount of work put in and the amount of energy 

produced is then the propulsive efficiency. In turbofans, a small amount of air is given a large increase in 

velocity to produce thrust. The higher the fan BPR, the larger the amount of air and the lower the velocity 

increase needed to produce a specific amount of thrust. In addition, higher bypass ratios mean lower fan 

pressure ratios and larger fan diameters; the fan blades do not have to work as hard to produce the

increase in velocity because there is more fan blade airfoil area doing the work, and so propulsive 

efficiency is higher. Open Rotors are the ultimate in high bypass ratio propulsion, nominally with BPRs 

between 40 and 80, and very low FPRs below 1.1. As seen in Figure 3, they have very low FPRs 

compared with other types of aircraft propulsion and thus very high values of propulsive efficiency. The 

potential gain in reduced fuel burn compared to other engine types is illustrated in Figure 8. General 

Electric Aviation (GE) has predicted that their Open Rotor concept has the potential to reduce fuel burn 

by as much as 25 to 30 percent compared to a current technology turbofan engines, the GE CFM56-5B. 

Figure 8.—Performance benefits of Open Rotor propulsion technology compared with 

current turbofan engines.
NASA/TM—2013-217690 8 

Figure 9.—Scale model Open Rotor fan model undergoing testing in the NASA Glenn 9- by15-Foot low 

speed (left) and 8- by 6-Foot high speed (right) wind tunnels.

Today, with the renewed emphasis on reducing the environmental impact of commercial aircraft, 

NASA and GE have partnered once again to investigate Open Rotor propulsion for the ERA N+2

generation of aircraft systems. GE has built upon its 1980s experience with the GE36 Program Unducted 

Fan (UDF), and has taken advantage of the tremendous increases in aerodynamic design capability in 

computational fluid dynamics, advanced high speed computers, and advanced strong light weight 

materials technology since the 1980s to develop a new generation of Open Rotors. GE is designing 

advanced three dimensional fan blade designs and using tailored material construction techniques to 

optimize the fan blade shapes for maximum performance, as well as minimum noise. An extensive series 

of tests of a first generation of advanced fan blade technology, sponsored by the SFW and ERA Projects, 

was conducted at NASA Glenn in 2010. Low speed testing up to Mach number 0.22 was performed in the 

Glenn 9- by 15-Foot wind tunnel on several fan blade designs to investigate their low speed performance 

near take-off conditions and at angle of attack. In Figure 9 (left side) a photo of the NASA/GE Open 

Rotor Propulsion Rig with pylon simulator installed is shown in the 9- by 15-Foot test section. Flow field 

diagnostics measurements using Phased Array, Particle Image Velocimetry and Pressure/Temperature 

Sensitive Paint were also made (Refs. 20 and 21) to investigate the interaction of the fan blade rows with 

each other and with a simulated pylon to characterize installation effects. Currently, NASA and GE are 

investigating the high speed cruise performance of Open Rotors up to Mach number 0.85 in the Glenn 

8’x6’ high speed wind tunnel. A photo of the NASA Open Rotor Drive Rig and GE Historical Baseline

blade set is shown on the right hand side in Figure 9. 

Preliminary test results show that GE has successfully demonstrated both aero performance and noise 

reduction objectives for this first generation of fan blade design (Refs. 22 and 23). Preliminary wind 

tunnel results have validated GE performance predictions of the high propulsive efficiency that can be 

achieved with Open Rotor propulsion and advanced 3D fan blade designs. Future NASA/GE 

collaborative testing of a second generation of Open Rotor fan blade designs at NASA Glenn, in 

partnership with ERA and the Federal Aviation Administration (FAA) Continuous Low Energy Efficient 

Noise (CLEEN) Program, is planned for mid 2011 to achieve further increases in performance as well as

reduce the noise signature of the Open Rotor system toward comparable turbofan levels. Figure 7 shows 

in the upper right hand corner the ultimate goal of the GE Open Rotor program for fuel burn and noise 

reductions. 
NASA/TM—2013-217690 9 

Engine Core Technology 

In the core turbomachinery area, the emphasis within the SFW and ERA Projects to reduce fuel burn 

is on increasing the power density of the engine core resulting in smaller units for a given thrust 

requirement (Refs. 24 and 25). Increasing the core compression system overall pressure ratio (OPR) while 

maintaining or improving aerodynamic efficiency is one area of focus. Another area is increasing the 

turbine inlet temperature (T4) to enable improved thermal efficiency. NASA is addressing the challenges 

of a higher OPR and higher T4 through a combination of developing advanced aerodynamic component 

designs, higher strength and higher temperature materials, and improved computational analysis

techniques. 

In Figure 11, the relationship between core thermal, propulsive and overall efficiencies are shown, 

along with the historic fuel burn trend with engine cycle and the goal for advanced engine technologies 

including the Ultra High Bypass ratio engine cycle. For the compression system, increasing the 

compressor OPR either drives the design toward more stages or a higher loading per stage. More stages 

mean a larger engine and more engine weight however. In either case, the overall efficiency of the 

compression system tends to suffer due to increased wetted area and drag losses or increased boundary 

layer separation and mixing losses, respectively. Overall engine size constraints and rotor dynamics issues 

limit the use of additional compressor stages, so often the solution to increased OPR is higher 

aerodynamic loading per stage. In Figure 11, the impact of overall pressure ratio on thermal efficiency is 

shown. For a given power requirement, higher stage loading can also enable a smaller core, since fewer 

compressor stages are required. 

Figure 10.—Trends in efficiencies with core and propulsor improvements.

Core 

Improvements

(direct impact on LTO NOx)

Propulsor

improvements
NASA/TM—2013-217690 10

Figure 11.—Trend in Brayton cycle thermal efficiency with compressor pressure ratio.

Figure 12.—Trend in specific core power gains with higher turbine inlet temperature (T4).

In Figure 12, the relationship between T4 and the specific power density of the core is shown, derived 

empirically from previous engine designs. While not having a first-order effect on thermal efficiency, 

raising T4 increases the thrust-to-weight of the engine. As the thermal efficiency and OPR increase, the 

compressor exit temperature (combustor inlet temperature), T3, increases due to increased compressive 

heating. For a fixed T4, the amount of allowable energy addition in the combustor decreases and the 

thrust of the engine must decrease for a given engine core flow rate. Therefore, increasing allowable T4 

enables engines having acceptable thrust-to-weight and core power density. For a given thrust and power 

requirement, then, higher T4 can allow smaller core components and overall engine size for a given BPR.

Current 

Technology

Current 

Technology

N+3 Goal
NASA/TM—2013-217690 11

This is important since the size of the engine increases to large diameters as BPR ratio increases (Figure 

3), resulting in unacceptable high engine weight and aircraft installation challenges. 

NASA and its research partners are investigating technologies to enable the high power density cores 

through advanced aerodynamic designs and the use of advanced, light weight and stronger materials for 

core components. In the aero design area (Figure 13), the emphasis of NASA research is being placed on 

improved design methods for high pressure, highly loaded compressors (Ref. 26) and turbines, novel 

turbine cooling flow techniques (Refs. 27 and 28) to improve cooling effectiveness, and novel flow 

control technologies for improved efficiency on compressor and turbine blades. As part of the ERA 

Project, NASA and its partner GE will be testing an advanced, multi stage, high speed, highly loaded 

compressor in the NASA Glenn W7 facility in early 2012. Detailed measurement of surface and flowfield 

quantities will be obtained in an effort to validate computational fluid dynamic (CFD) codes as well as to 

understand the physics of compressor blade row interaction and stage matching in a high speed, highlyloaded, close-coupled environment. This will be a benchmark study for the community to validate tools 

and improve highly-loaded compressor designs. It is expected that three-dimensional blade design 

methods and designs accounting for the blade row interaction effects will be incorporated into the 

advanced compressor design, resulting in a higher TRL through this experimental effort. NASA multistage turbomachinery codes will also be validated and extended through this research to account for these 

inter-blade row effects. These improvements will then be available to the broader gas turbine engine 

community for the advancement of highly-loaded turbomachinery designs and reduced fuel burn engines. 

In the materials area (Figure 14), NASA and its partners are investigating methods to improve cycle 

efficiency by improving turbine seals; improving thermal barrier coatings on turbine blades to reduce 

cooling load and improve thermal efficiency; developing high strength, high temperature metal alloys to 

allow higher blade loadings; developing shaped memory alloys to enable varying the engine cycle for 

improved efficiency; and high temperature ceramic matrix composites (CMC) for turbine vane and engine 

nozzle components to allow for higher engine core temperatures, reduce cooling flow requirements, and 

lower component weight. Currently, the ERA Project is conducting research to advance the TRL of CMC 

components through design, fabrication and demonstration testing of larger, more complex models in a 

relevant environment at NASA and partner facilities. 

Figure 13.—Engine core aero design-based technologies enabling advanced engine cycles.

Highly-Loaded, Multistage 

Compressor (higher 

efficiency and operability)

Low-Shock Design, 

High Efficiency, 

High Pressure 

Turbine

Aspiration Flow Controlled, 

Highly-Loaded, 

Low Pressure Turbine

Compressor 

Synthetic Jet Flow 

Control (reduced 

flow losses)

Low Pressure 

Turbine Plasma 

Flow Control

FLOW

ωz

UW

FLOW

Novel Turbine Cooling Concepts
NASA/TM—2013-217690 12

Figure 14.—Engine core materials-based technologies enabling advanced engine cycles.

Ceramic-matrix composites (CMC) offer opportunities for revolutionary changes in propulsion 

system design and operation. The lower density and higher temperature capability of CMC components, 

relative to that of metallic components, offer multiple engine advantages, such as weight saving, 

efficiency and thrust improvements, and reduced specific fuel consumption. Figure 15 shows areas within 

a typical aircraft engine core where CMC technology can be applied, design characteristics, and potential 

benefits to the engine. NASA has been developing durable, high-temperature CMCs with silicon carbide 

(SiC) matrices and SiC or carbon fibers for high-temperature structural applications (Ref. 29). These 

SiC/SiC composites are designed to be lightweight (~30 percent of metal density). With the advances in 

the development of thermal/environmental barrier coatings (TEBCs) for CMC applications, also under the 

ERA program, the performance of CMC components has improved significantly. Recently, SiC/SiC CMC 

combustor liner and turbine vanes have demonstrated cyclic durability at temperatures up to 1650 °C 

(3002 °F) (Ref. 30), a temperature significantly higher than the thermal capability of metal alloys 

(~1093 °C or ~2000 °F). A previous NASA program demonstrated the capability to fabricate complex 

turbine airfoil shapes from SiC/SiC composites that had increased operating temperature capability 

(2400 °F) and improved long-term durability. A representative SiC/SiC turbine airfoil cross section, 

shown in Figure 16, demonstrated the capability that was developed during that program to fabricate the 

complex features needed for turbine vanes. The U.S. aerospace industry, GE Aviation in particular, has 

since advanced this capability to the point that prototype SiC/SiC low pressure turbine vanes are currently 

under evaluation.

CMC material technology is also being investigated for high pressure turbine vanes. By reducing the 

cooling flow requirements for the high pressure turbine vanes, the overall efficiency of the core increases 

since more energy is available for extraction by the turbine sections. NASA is teamed with Rolls Royce 

Liberty Works to develop CMC technology for exhaust mixer nozzles. Material structural integrity is 

being demonstrated through testing of small material samples, and nozzle performance is being validated 

using subscale nozzle designs made of CMC material. ERA is also conducting research on CMC based 

turbine vanes. Fabrication technology is being developed to produce CMC turbine vanes without internal 

Low Leakage, Non-Contacting, Brush/Figure 

Turbine Seal (higher efficiency)

50 μm

Undoped layer

Eu-doped layer

Tb-doped layer

Bond coat 50 μm

Undoped layer

Eu-doped layer

Tb-doped layer

Bond coat

Multifunctional Thermal 

Barrier Coatings (higher 

temperature capability)

Variable Area Fan 

Nozzle using Shaped 

Memory Alloy (enables 

UHB engine cycle)

Multifunctional Structure with Power Harvesting 

Capability (increased thermal efficiency)

Porous,

Cellular

Structure

Thermoelectric legs

Piezoceramic 

Actuation Active Fan 

Blade Vibration 

Control (enhanced 

distortion tolerance)

Superalloy Materials based 

Turbomachinery Disk (higher 

temperature capability)
NASA/TM—2013-217690 13

material defects and with the same design features required for conventional material vane designs, 

including imbedded air cooling holes and thermal barrier coatings for high temperature applications. 

Figure 15.—Engine component applications for CMC materials and potential benefits.

Figure 16.—Generic low pressure turbine 

vane airfoil shape made with CMC 

SiC/SiC material.
NASA/TM—2013-217690 14

Figure 17.—Predicted reduction in Thrust Specific Fuel Consumption 

with reduced cooling load with CMC High Pressure Turbine Vanes.

A recent system study (Ref. 31) on an ultra-high bypass ratio (UHB) turbofan engine for an advanced 

twin-engine single-aisle transport (B737/A320 class aircraft) has shown that CMC turbine vanes 

(assuming 3000 °F temperature capability) enable the reduction of turbine cooling and therefore increases 

the turbine efficiency, which enabling an average TSFC reduction of 1.5 percent for an engine with 

Overall Pressure Ration of 32. The TSFC benefit increases with OPR as shown in Figure 17. The study 

also showed that CMC materials are an enabling technology for increased bypass ratio, and thus better 

propulsive efficiency. Higher temperature materials allow increased turbine rotor inlet temperature, which 

increases specific power of the engine, allowing a higher bypass ratio. Also, as Tbulk is allowed to 

increase, less turbine chargeable cooling is required. Core size shrinks as the demand of cooling air 

decreases. 

Summary

The NASA Aeronautics Research Mission Directorate has established two Projects to investigate,

develop, and demonstrate in partnership with U.S. Industry, Academia, and other Government Agencies, 

advanced subsonic transport aircraft technology to significantly reduce the impact of aircraft on the 

environment. The Subsonic Fixed Wing Project, under the Fundamental Aeronautics Program, 

established in 2006, is tasked to support fundamental research to Technology Readiness Level of 1 to 3 

for application to aircraft designs beyond 2025, or N+3 generation of aircraft. The Environmentally 

Responsible Aviation Project, under the Integrated Systems Research Program, is tasked to further 

develop the aircraft technology showing potential for meeting the system environmental goals to a 

Technology Readiness Level of 4 to 6 by 2020, or N+2 generation aircraft. 

NASA and its partners under the SFW and ERA Projects to date have demonstrated significant 

progress toward meeting the NASA Subsonic Transport System Level Metrics for installed fuel burn 

reduction with: advanced engine technology research in the areas of Ultra High Bypass and Open Rotor 

engine cycles to increase propulsive efficiency; advanced engine core research to produce higher 

propulsive and thermal efficiencies with higher Overall Pressure Ratio and Turbine Inlet Temperatures; 

advanced, three dimensional aerodynamic compressor and turbine designs; and advanced materials 

research in Ceramic Metal Composites and Superalloys to allow higher blade loading per stage to reduce 

engine weight as well as reduce turbine cooling requirements to increase overall efficiency. Project 

research includes computational design and analysis, including code validation; subscale model testing to 

verify technologies in laboratory and in relevant, complex, operational environments; and finally system 

level demonstrations of full scale components. 
NASA/TM—2013-217690 15

Appendix.—Definition of Technology Readiness Levels

TRL 1 Basic principles observed and reported: Transition from scientific research to applied research. 

Essential characteristics and behaviors of systems and architectures. Descriptive tools are mathematical

formulations or algorithms. 

TRL 2 Technology concept and/or application formulated: Applied research. Theory and scientific 

principles are focused on specific application area to define the concept. Characteristics of the application 

are described. Analytical tools are developed for simulation or analysis of the application.

TRL 3 Analytical and experimental critical function and/or characteristic proof-of-concept: Proof

of concept validation. Active Research and Development (R&D) is initiated with analytical and

laboratory studies. Demonstration of technical feasibility using breadboard or brassboard 

implementations that are exercised with representative data.

TRL 4 Component/subsystem validation in laboratory environment: Standalone prototyping 

implementation and test. Integration of technology elements. Experiments with full-scale problems or 

data sets. 

TRL 5 System/subsystem/component validation in relevant environment: Thorough testing of 

prototyping in representative environment. Basic technology elements integrated with reasonably 

realistic supporting elements. Prototyping implementations conform to target environment and interfaces. 

TRL 6 System/subsystem model or prototyping demonstration in a relevant end-to-end 

environment (ground or space): Prototyping implementations on full-scale realistic problems. Partially

integrated with existing systems. Limited documentation available. Engineering feasibility fully

demonstrated in actual system application.

TRL 7 System prototyping demonstration in an operational environment (ground or space): System

prototyping demonstration in operational environment. System is at or near scale of the operational

system, with most functions available for demonstration and test. Well integrated with collateral and

ancillary systems. Limited documentation available.

TRL 8 Actual system completed and "mission qualified" through test and demonstration in an 

operational environment (ground or space): End of system development. Fully integrated with

operational hardware and software systems. Most user documentation, training documentation, and

maintenance documentation completed. All functionality tested in simulated and operational scenarios. 

Verification and Validation (V&V) completed. 

TRL 9 Actual system "mission proven" through successful mission operations (ground or space): 

Fully integrated with operational hardware/software systems. Actual system has been thoroughly 

demonstrated and tested in its operational environment. All documentation completed. Successful

operational experience. Sustaining engineering support in place.
NASA/TM—2013-217690 16

References

1. National Aeronautics Research and Development Policy and Plan, 

http://www.ostp.gov/cs/nstc/documents_reports 

2. Madavan, Nateri, “NASA Subsonic Fixed Wing (SFW) Project Research Overview,” presentation at 

the ASME Turbo Expo 2011, Jun 2011. 

3. Hughes, Christopher E., “An Overview of the NASA Fundamental Aeronautics Program Subsonic 

Fixed Wing Project and Ultra High Bypass Partnership Research Goals,” presentation at the 15th 

AIAA/CEAS Aeroacoustics Conference (30th AIAA Aeroacoustics Conference), May 2009. 

4. Collier, Fayette; “Environmentally Responsible Aviation - Real solutions for environmental 

challenges facing aviation,” Report NF1676L-11223 (URL: http://ntrs.nasa.gov), GARDN, First 

Green Aviation Conference, Feb 2011. 

5. Berton, J., and Guynn, M.: “Multi-Objective Optimization of Turbofan Design Parameters for an 

Advanced, Single-Aisle Transport,” AIAA Paper 2010-9168, AIAA 10th Aviation Technology, 

Integration, and Operations (ATIO) Conference, Sep 2010. 

6. Guynn, M.; Berton, J.; Fisher, K.; Haller, W.; Tong, M.; and Thurman, D.: “Refined Exploration of 

Turbofan Design Options for an Advanced Single-Aisle Transport Aircraft,” NASA/TM—2011-

216883, Jan2011.

7. Tong, Michael T.; Jones, Scott M.; Haller, William J,; Handschuh, Robert F.; “Engine Conceptual 

Design Studies for a Hybrid Wing Body Aircraft,” NASA/TM—2009-215680, Nov 2009. 

8. Hughes, Christopher, NASA Glenn Research Center; Schweitzer, Jeff, Pratt & Whitney; “Ultra High 

Bypass Ratio Engine Research for Reducing Noise, Emissions, and Fuel Consumption,” Document 

20080005581 (URL: http://ntrs.nasa.gov), NASA Fundamental Aeronautics Program 1st Annual 

Review, Oct 2007. 

9. Hughes, Christopher E., “Collaborative Research on the Ultra High Bypass Ratio Engine Cycle to 

Reduce Noise, Emissions and Fuel Consumption,” Report E-16711 (URL: http://ntrs.nasa.gov), 

UTIAS-MITACS International Workshop on Aviation and Climate Change, May 2008. 

10. Hughes, Christopher E., “NASA Partnerships and Collaborative Research on Ultra High Bypass 

Cycle Propulsion Concepts,” Report E-16904 (URL: http://ntrs.nasa.gov), Acoustics Technical 

Working Group Meeting, Sept 2008. 

11. Hughes, Christopher E., “NASA Collaborative Research on the Ultra High Bypass Engine Cycle and 

Potential Benefits for Noise, Performance and Emissions,” ISABE-2009-1274, 2009. 

12. Hughes, Christopher E., NASA Glenn Research Center; Lord, Wesley, Pratt & Whitney; 

“NASA/Pratt & Whitney Collaborative Partnership Research in Ultra High Bypass Cycle Propulsion 

Concepts,” Report E-16905 (URL: http://ntrs.nasa.gov), Fundamental Aeronautics Program 2nd 

Annual Review, Oct 2008. 

13. Hughes, Christopher, “Geared Turbofan Technology,” presentation at the NASA Green Aviation 

Summit, September 2010, URL: www.aeronautics.nasa.gov/pdf/hughes_green_aviation_summit.pdf. 

14. Hughes, Christopher, and Smith, Steven; “The Promise and Challenges of Ultra High Bypass Ratio 

Engine Technology and Integration,” Report HQ-STI-11-012 (URL: http://ntrs.nasa.gov), AIAA 49th

Aerospace Sciences Meeting, Jan 2011.

15. Hughes, C., and Edmane, E., “Overview of Recent Ultra High Bypass Engine Cycle-Based Scale 

Model Test Results,” presentation at the Fundamental Aeronautics Program 3rd Annual Review, Sep 

2009. 

16. Hughes, Christopher, NASA Glenn Research Center; Zeug, Theresa, GE Aviation; “NASA/GE 

Aviation Collaborative Partnership Research in Ultra High Bypass Cycle Propulsion Concepts,” 

Report E-16903 (URL: http://ntrs.nasa.gov), Fundamental Aeronautics Program 2nd Annual 

Review, Oct 2008. 

17. Van Zante, Dale, and Breeze-Stringfellow, Andrew, “The NASA/GE Collaboration on Open Rotor 

Testing,” presentation at the University of Toronto Institute for Aerospace Studies: Workshop on 

Aviation and Climate Change, May 2010. 
NASA/TM—2013-217690 17

18. Van Zante, Dale, and Breeze-Stringfellow, Andy, “The NASA/GE Collaboration on Open Rotor 

Testing,” presentation at the Royal Aeronautical Society, Conference on Applied Aerodynamics: 

Capabilities and Future Requirements, Jul 2010. 

19. Hendricks, Eric, “Development of an Open Rotor Cycle Model in NPSS using s Multi-Design Point 

Approach,” GT2011-46694, ASME Turbo Expo 2011, Jun 2011. 

20. Van Zante, Dale E., “NASA/GE Collaboration on Open Rotors - High Speed Testing,” Report E17798 (URL: http://ntrs.nasa.gov), Acoustics Technical Working Group, NASA Glenn Research 

Center, May 2011.

21. Van Zante, Dale, “Reestablishing Open Rotor as an Option for Significant Fuel Burn 

Improvements,” Report HQ-STI-11-013 (URL: http://ntrs.nasa.gov), AIAA 49th Aero Sciences 

Meeting, Jan 2011.

22. Van Zante, Dale; Wojno, John; “The NASA/GE Open Rotor Research Campaign: ERA Diagnostics 

Test,” presentation at the ASME Turbo Expo 2011, Jun 2011. 

23. Van Zante, Dale; Gazzaniga, John; Elliott, David; and Woodward, Richard; “An Open Rotor Test 

Case: F31/A31 Historical Baseline Blade Set,” ISABE-2011-1310, Sep 2011. 

24. Hiedmann, James, “Advanced Core Engine Technology,” presentation at the NASA Green Aviation 

Summit, Sep 2010, URL: www.aeronautics.nasa.gov/pdf/hughes_green_aviation_summit.pdf.

25. Heidmann, James; “Improving Engine Efficiency through Core Developments,” presentation at the 

AIAA 49th Aero Sciences Meeting, Jan 2011.

26. Lurie, E. A., Van Slooten, P.R., Medic, G., Mulugeta, J.M., Holley, B. M., Feng, J., Sharma, O., and 

Ni, R., “Design of a High Efficiency Compact Centrifugal Compressor for Rotorcraft Applications,” 

Proceedings of AHS Forum 67, May 2011. 

27. Heidmann, J., and Ekkad, S.V., A Novel Anti-Vortex Turbine Film Cooling Hole Concept, ASME 

Journal of Turbomachinery, Jul 2008, Vol. 130. 

28. Dhungel, A., Lu, Y., Phillips, W.A., Ekkad, S.V., Heidmann, J., Film Cooling from a Row of Holes 

Supplemented with Anti-vortex Holes, ASME Journal of Turbomachinery, Jan 2009, Vol. 131. 

29. DiCarlo, J. A., et al, “High-Performance SiC/SiC Ceramic Composite Systems Developed for 1315 

°C (2400 °F) Engine Components,” NASA/TM—2004-212729, 2004, pp.12-13. 

30. Zhu, D., Miller, R. A., Fox, D. S., “Thermal and Environmental Barrier Coating Development for 

Advanced Propulsion Engine Systems,” NASA/TM—2008-215040. 

31. Tong, M. T., “An Assessment of the Impact of Emerging High-Temperature Materials on Engine 

Cycle Performance,” GT2010-22361, ASME Turbo Expo 2010, Jun 2010. 
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The NASA Fundamental Aeronautics Program Subsonic Fixed Wing Project and Integrated Systems Research Program Environmentally 

Responsible Aviation Project in the Aeronautics Research Mission Directorate are conducting research on advanced aircraft technology to 

address the environmental goals of reducing fuel burn, noise and NOx emissions for aircraft in 2020 and beyond. Both Projects, in 

collaborative partnerships with U.S. Industry, Academia, and other Government Agencies, have made significant progress toward reaching 

the N+2 (2020) and N+3 (beyond 2025) installed fuel burn goals by fundamental aircraft engine technology development, subscale 

component experimental investigations, full scale integrated systems validation testing, and development validation of state of the art 

computation design and analysis codes. Specific areas of propulsion technology research are discussed and progress to date.

15. SUBJECT TERMS

Ultra high bypass; ERA; Wind tunnel 

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