Michael D. Hathaway and Ruben Del Rosario

Glenn Research Center, Cleveland, Ohio

Nateri K. Madavan

Ames Research Center, Moffett Field, California

NASA Fixed Wing Project Propulsion Research and

Technology Development Activities to Reduce

Thrust Specifi c Energy Consumption

NASA/TM—2013-216548

July 2013
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Michael D. Hathaway and Ruben Del Rosario

Glenn Research Center, Cleveland, Ohio

Nateri K. Madavan

Ames Research Center, Moffett Field, California

NASA Fixed Wing Project Propulsion Research and

Technology Development Activities to Reduce

Thrust Specifi c Energy Consumption

NASA/TM—2013-216548

July 2013

National Aeronautics and

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Cleveland, Ohio 44135

Prepared for the

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San Jose, California, July 14–17, 2013
Acknowledgments

This effort is sponsored by the FW project of the NASA Fundamental Aeronautics Program. The research described is part of the 

FW Propulsion Effi ciency subproject portfolio. The authors greatly appreciate the accomplishments of the many researchers and 

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This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2013-216548 1 

NASA Fixed Wing Project Propulsion Research and 

Technology Development Activities to Reduce 

Thrust Specific Energy Consumption 

Michael D. Hathaway and Ruben Del Rosario 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Nateri K. Madavan 

National Aeronautics and Space Administration 

Ames Research Center 

Moffett Field, California 94035 

Abstract 

This paper presents an overview of the propulsion research and technology portfolio of NASA 

Fundamental Aeronautics Program Fixed Wing Project. The research is aimed at significantly reducing the 

thrust specific fuel/energy consumption of notional advanced fixed wing aircraft (by 60 percent relative to a 

baseline Boeing 737-800 aircraft with CFM56-7B engines) in the 2030 to 2035 time frame. The research 

investments described herein are aimed at improving propulsive efficiency through higher bypass ratio fans, 

improving thermal efficiency through compact high overall pressure ratio gas generators, and exploring the 

potential benefits of boundary layer ingestion propulsion and hybrid gas-electric propulsion concepts. 

1.0 Introduction 

In 2008, the NASA Fixed Wing (FW) Project (then known as the Subsonic Fixed Wing project) 

funded system studies (Refs. 1 to 5) of notional future fixed-wing aircraft concepts that could be viable 

for commercial aviation in the 2030 to 2035 time frame. NASA required that these concept aircraft meet 

very aggressive metrics for fuel/energy consumption, emissions, and noise. A detailed description and 

assessment of propulsion systems and technologies considered in these system studies, and recommendations for technologies that merited further investigation, were reported by Ashcroft, et al. Reference 6. 

The FW Project has identified several propulsion system themes that have emerged from the system 

studies as showing potential to significantly reduce fuel/energy consumption. These propulsion system 

themes are (1) cleaner, compact higher bypass ratio propulsion, (2) unconventional propulsion-airframe 

integration, and (3) hybrid gas-electric propulsion. The FW project has established the following highlevel propulsion system technical challenges based on these themes to guide research investments; (1) 

compact high overall pressure ratio (50+ OPR) gas generators to increase fan bypass ratio and improve 

propulsive and thermal efficiency, (2) integrated boundary layer ingestion system to achieve a vehiclelevel net system fuel/energy consumption benefit on a representative vehicle, by improving propulsive 

efficiency and reducing aircraft drag, and (3) high power density electric motors (four times existing 

state-of-the-art, SOA), and power management and distribution for future notional hybrid gas-electric 

propulsion systems. These investments are aimed at meeting the challenging metric of reducing thrust 

specific fuel/energy consumption by 60 percent relative to a baseline Boeing 737-800 aircraft with 

CFM56-7B engines. 

 
NASA/TM—2013-216548 2 

The specific propulsion system research and development activities currently funded under the FW 

Project to address these high level technical challenges are described in this paper. These activities 

include research conducted at the four NASA Aeronautics Research Centers (Glenn, Langley, Ames, and 

Dryden), as well as work done under contracts and cooperative agreements by academic and industry 

partners to augment NASA in-house efforts. 

2.0 Propulsion System Technical Challenges and Research Portfolio 

As shown in Figure 1, improvements in both gas turbine engine cores (i.e., gas generators) and 

propulsors (i.e., fans) contribute to reducing thrust specific fuel consumption (TSFC). Advancements in 

core and propulsor technologies to date have led to high bypass ratio (BPR) engines, such as those 

deployed on Boeing 747/777 aircraft, that have provided significant reductions in thrust specific energy 

consumption relative to the original Whittle engine and turbojet engines. Further reductions in thrust 

specific fuel consumption are possible with technology advancements to enable ultra-high BPR fans and 

engine cores operating at higher overall pressure ratio (OPR) for increased propulsive and thermal 

efficiency. Given the physical constraints on maximum fan nacelle diameter, an alternative approach to 

increasing fan BPR is to reduce the engine core diameters and flow size. In order for a smaller flow size 

core to provide the power necessary to drive the higher bypass ratio fan the engine OPR must increase to 

around 50+; the higher OPR increases the engine thermal efficiency as shown in Figure 2, and contributes 

to reducing TSFC. It is also necessary to increase the temperature of the air entering the turbine section 

that extracts energy from the air to drive both the core and the fan. The more energy that can be put into 

the air, raising its temperature, the more energy that is available for the turbine to extract. If more energy 

is available for extraction, the turbine size can be reduced, leading to a smaller lighter engine and less fuel 

expended for the aircraft to carry it. Current turbine inlet temperature technology limits and the goal of 

the FW project 2030 to 2035 time frame for “N+3” generation engines are shown in Figure 3. Increases in 

the turbine inlet temperature will require either advanced materials that can withstand the higher 

temperatures, or cooling the turbine blades with high-pressure air from the exit of the compressor. The 

latter option reduces the overall efficiency of the engine by diverting high-pressure air from the 

combustor and expending energy in routing air from the compressor to supply the turbine cooling. 

The higher 50+ OPR engine cycle also results in higher temperatures in the aft core stages, requiring 

advances in blade and disk material capability for maximum turbomachinery component efficiency and in 

order to reduce or eliminate chargeable turbine cooling air. The 50+ OPR engine results in compressor 

exit temperatures on the order of 1500 F (816 C) near the rotor rim, which is the same air used for 

turbine bore cooling purge flow, and exceeds current turbine disk material capability. Also, in order to 

accommodate the HPT temperatures commensurate with the 50+ OPR engine. the first stage stator 

Figure 1.—Overall gas turbine engine efficiency as a 

function of core thermal efficiency, propulsive and 

transmission efficiencies.

Figure 2.—Thermal efficiency as a function of compressor 

overall pressure ratio (OPR). 
NASA/TM—2013-216548 3 

Figure 3.—Specific core power as a function of turbine inlet temperature. 

vanes and rotor blades must withstand surface temperatures on the order of 3000 F (1649 C), and 

2700 F (1482 C) with adequate, likely full-life, service capability. Technology investments are therefore 

required to develop compressor and turbine disk materials and high pressure seals that can accommodate 

1500 F (816 C) conditions, and ceramic matrix composite (CMC) vanes and blades that can accommodate 3000 F (1649 C) and 2700 F (1482 C) temperatures, respectively, without requiring blade cooling 

flow. A more in-depth assessment of aeropropulsion for commercial aviation in the 21st Century and 

research directions needed with respect to gas turbine propulsion is provided by Epstein (Ref. 8). 

The potential for reduced aircraft drag and possibly improved lift exists from propulsors that are 

highly integrated with the wing or fuselage and may serve as “flow control devices” by ingesting 

significant portions of the wing or aircraft boundary layer flow and filling in the aircraft wing or fuselage 

wake to reduce aircraft drag and potentially improve lift. The drawback of having propulsors that are 

highly integrated with the airframe or wing is the consequent detrimental impact on the propulsor (fan) 

performance and life from having to continually ingest a distorted inlet flow profile, thereby reducing the 

efficiency and stall range of the fan relative to operating with undistorted inflow conditions. To realize the 

maximum performance benefit for fuel burn reduction it is necessary to minimize the fan inflow 

distortion and design a robust fan stage tolerant to continuous ingestion of inlet distortion. A fan operating 

with continuous inlet distortion may have lower performance potential (peak efficiency and stall range) 

relative to one designed to operate continuously with undistorted inflow at cruise conditions. Technology 

investments are therefore required to assess the performance potential of boundary layer ingesting 

propulsion systems that are highly integrated with the aircraft wing or fuselage, and to determine the 

tolerable fan stage efficiency and stall margin decrement that still results in significant fuel burn reduction 

while also being structurally robust for full-life service capability. 

Finally, future hybrid gas-electric vehicle concepts hold promise in providing significant reductions in 

fuel/energy consumption, emissions, and noise. Investments in such far-term technologies are thus 

warranted, but these benefits will likely require major technical breakthroughs. The key “long pole” hybrid 

gas-electric technologies identified by the FW Project for investment are the development of electric motors 

and components with four times the power density of the current SOA, and the management and distribution 

of megawatt-range electric power in the aircraft for the duration of its flight mission. 

 
NASA/TM—2013-216548 4 

As detailed in the discussion above, the FW Project is currently investing in research with the 

potential to contribute to overcoming the identified high-level propulsion system technical challenges. 

Such research includes: 

1. Enabling higher temperature capable disk and turbine blade/stator materials. 

2. Reducing endwall and tip leakage losses of low exit corrected flow gas generator stages with 

large tip clearance gap to accommodate the high temperatures and retain thermal efficiency 

benefits of 50+ OPR gas generators. 

3. Enabling efficient, robust high bypass ratio fans tolerant to inflow distortion from highlyintegrated propulsion systems that ingest aircraft boundary layer flow. 

4. Enabling efficient higher power density electric motors for future hybrid gas-electric propulsion 

aircraft concepts. 

In addition to the above, other research and technology development activities of interest include: 

enabling capability for managing and distributing power (eventually up to 50 MW) for future hybrid gaselectric propulsion concepts, and developing fan blades that adapt to engine operating conditions 

throughout the aircraft flight envelope. Of course, there are a number of other potential technologies to 

enable advancements in future commercial aviation transports, but limited resources make it necessary to 

prioritize and focus on a limited set of promising technologies with the highest potential impact and with 

the perceived ability to be realized in the 2030 to 2035 time frame. The research and technology 

development activities currently funded by the FW Project are discussed according to the various 

propulsion research themes described above in the following sections. 

2.1 Cleaner, Compact High Bypass Ratio Propulsion 

The “cleaner” aspect of this high level challenge pertains to the reduced emissions goal obtained from 

combustor technology investments and is therefore not described in this paper. As stated above, the 

technical challenges to achieving compact high bypass ratio propulsion systems are centered around two 

aspects: (1) achieving a smaller, compact, high efficiency gas generator to enable increased bypass ratio 

for a fixed diameter nacelle, thereby improving propulsive efficiency, and (2) increasing overall pressure 

ratio (OPR) of the gas generator, and hence improving thermal efficiency, to achieve the aircraft power 

and propulsion requirements with a smaller flow size gas generator while also operating with uncooled 

turbine blades to reduce chargeable cooing. The former entails mitigating the losses associated with the 

larger tip clearance gap leakage and endwall flows, and the latter requires development of high 

temperature materials and seals to enable the high OPR engine. 

2.1.1 Mitigating Tip Clearance Gap Leakage and Endwall Loss 

The smaller, compact, high efficiency gas generator operating at high OPR (50+) results in aft stages 

operating with very low, on the order of 1 to 2 lbm/s (0.45 to 1 kg/s), exit corrected flows with blade 

heights on the order of 0.5 in. (1.27 cm) or less and blade tip-clearance gaps on the order of 4 percent of 

blade span, relative to current state-of-the-art gas generators with exit corrected flows on the order of 

6 lbm/s (2.7 kg/s), blade heights of 0.75 in. (1.91 cm) and blade tip-clearance gaps on the order of 1 to 

2 percent of blade span. The resultant larger blade tip-clearance gaps increase endwall/tip clearance gap 

leakage losses, reducing component efficiency (thereby reducing thermal efficiency benefits), and the 

small blades challenge manufacturing capability and blade robustness. 

To help address the tip clearance gap leakage and endwall losses the FW Project initiated a NASA 

Research Announcement (NRA) call for proposals to improve understanding and mitigate tip clearance 

gap leakage and end wall losses in high pressure ratio cores (Ref. 7). Five three-year NRAs were awarded 

in the summer of FY11 to teams from: Purdue University/Rolls-Royce, Johns Hopkins University (JHU), 

Pratt & Whitney/Penn State University (PSU), Honeywell/University of Notre Dame (UND), and U.S. 

Naval Academy/Cleveland State University (CSU). All the NRAs involve detailed measurements and 
NASA/TM—2013-216548 5 

computational analyses for at least two tip clearance gaps to elucidate the fundamental physics of the 

endwall/tip clearance leakage flows, and some of them propose to test concepts to mitigate losses and 

improve efficiency of small cores with large tip clearance gaps. Two of the NRAs focus on compressor 

and three on turbine endwall/tip clearance flows. The FW project also awarded a Phase II NRA to 

Massachusetts Institute of Technology (MIT)/Pratt and Whitney to study and identify the technical 

challenges to developing small core compressors and investigate novel concepts to address the small core 

challenge. 

Professor Katz and his students at JHU are addressing compressor tip clearance/endwall flows in a 

1-1/2 stage compressor with acrylic blades and casing and a working fluid that is a concentrated solution 

of sodium iodide (NaI) in water (63 percent by weight) to match the optical refractive index of the blades 

and casing material. The casing and blades thus become almost invisible, allowing unobstructed 

measurements almost everywhere without concern for flare light from the laser beams impinging on 

metallic surfaces. This technique provides the opportunity for much greater detailed measurements of the 

tip clearance/endwall flow field than is possible with metallic blading, especially as the blades cut through 

the path of the laser beam. The JHU 1-1/2 stage compressor is a reduced scale model of the upstream 

1-1/2 stages of the NASA Low Speed Axial Compressor (LSAC), which in turn was patterned after the 

GE Energy Efficency Engine (E3

) geometry. To retain adequate rotor blade thickness for structural 

integrity, the rotor blade tip was required to be 0.25 in. (6.35 mm) thick. Due in part to power limitations 

of the JHU facility, which pumps a mixture of water and Nal, a direct scale model of the LSAC geometry 

was not deemed feasible. In order to maintain solidity for aero-similitude, the rotor and stator chords were 

increased and the number of rotors and stators decreased. Numerical simulations of the scaled geometry 

showed satisfactory operation. Figure 4 is a schematic of the JHU turbomachinery facility with the new 

NASA compressor setup installed, and Figure 5 is an enlarged cross section of the new NASA 1-1/2 stage 

axial turbomachine. A blow-up of just the 1-1/2 stage axial turbomachine indicating transparent/acrylic 

Figure 4.—The JHU turbomachine facility with the new NASA compressor setup installed. 

Figure 5.—A magnified cross section of the complete NASA 1-1/2 stage axial turbomachine setup. 
NASA/TM—2013-216548 6 

hardware shown in Figures 6 and 7 includes a 3D representation of the three compressor blade rows and a 

photographic view from the inlet guide vane (IGV) inlet, within the transparent casing, showing the 

transparent/acrylic blades and hub with bearing assembly. The NaI working fluid with index of refraction 

matched to the transparent/acrylic components provides unobstructed optical access, even through the 

rotating blades, thereby facilitating Stereo PIV (partical image velocimetry) in axial and circumferential 

planes through the compressor, Figure 8, and Tomograpic PIV, Figure 9, which provides full 3D time 

resolved measurements to capture details of the leakage vortex. 

Figure 6.—Blow up of the 1-1/2 stage axial turbomachine depicting transparent/ 

acrylic hardware. 

Figure 7.—(a) Three compressor blade rows. Flow is from right to left; and (b) photographic view from 

IGV inlet, within the transparent casing, showing the acrylic blades and hub with bearing assembly.
NASA/TM—2013-216548 7 

 (a) (b) 

Figure 8.—Stereo PIV setup for (a) axial and (b) blade-to-blade planes. 

Figure 9.—Setup of tomographic PIV system. 

 
NASA/TM—2013-216548 8 

Professor Key and her students at Purdue University are testing a Rolls-Royce designed 3-1/2 stage 

compressor, Figures 10 and 11. The compressor will be tested in the Purdue 3-stage research facility with 

1.5, 3, and 4 percent tip clearance gaps based on annulus height. This allows comparison of typical tighter 

clearances achievable for larger core sizes with the larger clearance gaps expected of small, ~1.5 lbm/s 

(3.3 kg/s), core compressors with blade spans less than 0.5 in. (13 mm). The rotor tip clearance is varied 

from the baseline clearance by recessing the casing over the rotors to effect larger rotor blade tip gaps. 

The Purdue effort is focused on obtaining detailed measurements of the overall and stage performance of 

the compressor via flow field traverses upstream and downstream of blade rows using steady 

instrumentation probes and unsteady high response kulites and hot-wires. In addition, volumetric 

illuminating particle image velocimetry, Figure 12, and Tomographic image velocimetry, Figure 13, may 

also be employed to capture greater detail of the endwall/tip clearance flow physics. Complementary 

steady and unsteady CFD simulations are also being conducted for additional insight and will be validated 

by the experimental data being acquired. Testing with conventional probes, hot-films, and high response 

casing kulites has already begun and the initial results look promising. 

Professor Volino at the U.S. Naval Academy is investigating the unsteady endwall and tip gap flows 

in turbine blades using a low speed cascade facility, with rotating “wake” generator bars to simulate the 

periodic unsteady flow from the adjacent upstream blade row, as shown in Figures 14 and 15. The E3

rotor blade tip section and blade angles were selected to match the geometry and blade angle settings 

recently tested in the NASA CW-22 high-speed cascade facility. The low speed cascade tests are intended 

to document the unsteady response of the turbine rotor endwall flow, with and without tip gaps, to 

unsteady wakes from adjacent upstream rotating bars using Kiel probes, hot-wires, PIV, and other 

measurement techniques. The measurements are intended to improve understanding of the turbine tip 

leakage and endwall flow physics with increased rotor blade tip clearance (~4 to 5 percent of blade span), 

Figure 10.—Meridional view of the Purdue 3-1/2 stage research 

compressor. 

Figure 11.—Purdue 3-1/2 stage research 

compressor hardware and blade geometry. 

 

Figure 12.—Volumetric illuminating PIV. Figure 13.—Tomographic PIV setup. 

 
NASA/TM—2013-216548 9 

Figure 14.—Schematic of USNA turbine cascade and 

rotating bar arrangement. 

Figure 15.—Photo of turbine cascade blades and 

rotating bar installed in USNA test facility. 

(a) (b) 

Figure 16.—(a) PSU low speed cascade (LSC), and (b) axial flow turbine research facility (AFTRF). 

to help guide techniques for controlling tip flows using passive or active blowing techniques to reduce 

engine performance degradation with respect to tight tip clearances (~1 to 2 percent of blade span). To 

complement the experimental measurements, currently underway, Professor Mounir Ibrahim at Cleveland 

State University is performing high fidelity CFD simulations of the experimental configuration. 

The Pratt & Whitney/PSU NRA is aimed at understanding and reducing high pressure turbine 

endwall/rim-cavity secondary flow and tip clearance losses. The experimental investigations involve 

testing in PSU’s low speed cascade (LSC), Figure 16(a), and high speed axial flow turbine research 

facility (AFTRF), Figure 16(b), which will provide detailed, benchmark-quality aerodynamic and cooling 

effectiveness data to validate CFD design analysis and verify optimization approaches to minimize 

turbine endwall, rim cavity, and endwall clearance losses. The three times scale LSC geometry and 

AFTRF test hardware have been aerodynamically matched as shown in Figure 17, and the features of the 

interchangeable rim-cavity and instrumentation are shown in Figure 18. The figure shows a movable bluff 

body to investigate dependency and characteristics of the cavity purge flow to the downstream rotor 

potential flow field. The LSC test section is complete, and the AFTRF hardware has been assembled and 

instrumentation development is underway. Initial LSC tests have begun and the next steps are to fabricate 

the LSC bluff body, complete the LSC test matrix, and then complete the baseline AFTRF testing. 

The Honeywell/UND effort involves the design and subsequent testing of a new HPT turbine in the 

UND turbine test facility, Figure 19. The initial HPT stage design parameters are provided in Table 1. 

CFD tools and design of experiments (DOE) are employed to identify high-pressure turbine (HPT) design 

parameters that influence the blade tip leakage and endwall losses and erode overall performance, and to 

select blade tip parameters and endwall contouring for optimized HPT stage efficiency. The analyses 

include a variety of CFD techniques (RANS/URANS, harmonic methods, LES and/or DES). 
NASA/TM—2013-216548 10 

(a) (b) 

Figure 17.—(a) Comparison of PSU 3x scale LSC blade geometry to AFTRF blade, and (b) matching of aerodynamic 

parameters. 

Figure 18.—LSC interchangeable rim cavity geometry and instrumentation, 

and movable downstream bluff body and instrumentation. 

(a) UND test facility (b) New HPT rotor 

Figure 19.—(a) UND test facility, and (b) new HPT rotor. 
NASA/TM—2013-216548 11 

TABLE 1.—INITIAL UND HPT DESIGN PARAMETERS 

Figure 20.—Blade tip design modifications proposed to reduce leakage loss. 

Four different concepts were considered in the design: (1) unloading the tip by changing tip geometry 

or by reducing work level at the tip, (2) implementing barriers to tip flow such as squealer tips, winglets 

or fences, as shown in Figure 20, (3) examining the impact of airfoil static pressure loading on the tip 

leakage vortex, and (4) endwall contouring. The optimized HPT design obtained from the initial CFD and 

DOE results has reduced tip reaction, 10 percent higher rotor blade count, 20 percent larger tip section 

axial chord, and 0.7 percent higher efficiency compared to the baseline design. The optimized stage is 

being built for testing in the UND transonic test facility. The resulting data will be analyzed and 

additional high fidelity calculations will be performed. 

The MIT/P&W NRA is aimed at identifying the technical challenges to developing small low-exitcorrected-flow aft-stage core compressors, and investigating novel concepts to address the small core 

challenge (Ref. 10). Analyses of Reynolds number effects and performance assessment of tip clearance 

gap in high pressure compressors has focused on the D8.5 engine for the MIT Double Bubble boundary 

layer ingesting aircraft concept. A comparison of the notional size of an anticipated future small core 

(denoted N+3) relative to a CFM56-7 core (~22 klbf, 100 kN, thrust class) is shown in Figure 21 and 

illustrates the challenge of increasing OPR for this engine class. MIT’s analysis of Reynolds number 

effects, Figure 22, indicate the D8.5 compressor mid-stages achieve the lowest Reynolds numbers. MIT 

has also investigated the performance impact of three HPC conceptual scaling configurations, Figure 23. 

1. Pure “photo” scale—Takes a modern compressor with a hub-to-tip ratio of 0.93 at the last stage 

and scales it to a lower corrected flow, 

2. Shaft limited—Assumes the LP shaft radius sets the minimum radius. Mean radius is larger than 

from photographic scaling, and 

3. Shaft removed—Eliminates the LP shaft constraint and pulls in the flow path. Hub-to-tip ratio 

decreased to 0.85 for a 1.5 lbm/s (0.68 kg/s) core. 

MIT then assessed the compressor efficiency impacts according to the following five steps: (1) 

determine blade sizes, Re numbers, and tip clearances, (2) calculate Re number efficiency penalty of each 

stage, (3) calculate tip clearance efficiency penalty of each stage, (4) calculate compressor polytropic 

efficiency, (5) determine aircraft fuel burn with Transport Aircraft System OPTimization (TASOPT) 

(Ref. 9) a program for simultaneously optimizing the airframe, engine, and operating parameters of a 

wing and tube transport aircraft. The results are shown in Figure 24. 
NASA/TM—2013-216548 12 

Figure 21.—Scaling CFM56-7 to future (denoted N+3) engine size 

(Ref. 10). 

Figure 22.—Reynolds number in D8.5 HPC stages (Ref. 10). 

Figure 23.—Impact of different scaling dependent on shaft constraint (Ref. 10). 
NASA/TM—2013-216548 13 

Figure 24.—Fuel burn impact as a function of HPC 

polytropic efficiency and scaling options (Ref. 10). 

(a) (b) (c)

Figure 25.—(a) Schematic of a hybrid disk section, (b) creep resistance of LDS alloys required for 816 C (1500 F) 

disk rim temperatures and (c) 3D strain mapping of PM/SX joint indicating most deformation occurs in the single 

crystal region (Ref. 11). 

2.1.2 Developing High Temperature Materials, Coatings, and Seals to Enable the Higher 50+ 

OPR Engine. 

To enable a 50+ OPR engine, material advancements are required to accommodate the higher 

compressor exit temperatures (around 1500 F, 816 C, near the rotor rim), and is the same air used for 

turbine bore cooling purge flow, requiring development of higher temperature capable turbine disk 

materials. Furthermore, to accommodate the HPT temperatures commensurate with the 50+ OPR engine the 

first stage stator vanes and rotor blades must withstand surface temperatures on the order of 3000 F 

(1649 C), and 2700 F (1482 C) with adequate, likely full life, service capability. Technology investments 

are therefore required to achieve compressor and turbine disk materials and high pressure seals that can 

accommodate 1500 F (816 C) conditions, and ceramic matrix composite (CMC) vanes and blades that can 

accommodate 3000 F (1649 C) and 2700 F (1482 C) temperatures, respectively, without requiring blade 

cooling flow. Both the high temperature disk and CMC blade material will require coatings for environmental and thermal protection. The development of CMC turbine materials to accommodate the high temperature engine cycle is being managed and supported by the Aeronautical Sciences Project under the NASA 

Fundamental Aeronautics Program and will therefore not be discussed herein.

The FW project is currently investing in the development of hybrid-disk materials to withstand 

1500 F (816 C) turbine rim conditions and 1300 F (704 C) bore temperatures, Figure 25(a), and 

coatings to protect against hot-corrosion. Hurst (Ref. 11) has reported a recent comprehensive overview 

of materials and structures research focused upon propulsion applications within the FW Project. Recent 

test results by materials researchers at the NASA Glenn Research Center indicate that the right combination of optimal superalloys and microstructures at the bore/web and rim could enable a 1500 °F (816 C) 

hybrid disk to accommodate compressor exit temperatures for the 50+ OPR engine. Materials have been 

selected and initial mechanical properties generated for further study of this hybrid disk concept: a 15 μm 

grain size low-solvus, high-refractory (LSHR) superalloy bore/web powder metal superalloy and a new 

NASA low density single (LDS) crystal rim superalloy (LDS1101+Hf). 
NASA/TM—2013-216548 14 

Figure 26.—GRC non-contacting 

brush/finger turbine seal. 

Figure 27.—GRC high temperature, high speed 

turbine seal test facility (Ref. 12). 

These materials are being tailored to meet the structural requirements of a disk with increased maximum 

operating temperature of 1500 F (816 C) of future high 50+ OPR engines, compared to current 1300 F 

(700 C), Figure 25(b) and (c). This significantly higher operating temperature would enable the high 

50+ OPR engine and thereby could improve engine operating efficiency and reduce fuel consumption. 

The research and technology development activities to enable the 1500 F (816 C) capable hybrid disk 

involves not only the development of the desired materials property characteristics for the super alloy 

power metallurgy bore/web and the single crystal rim, but also the joining process, including transition 

location, and coatings for thermal and environmental protection. 

The FW project is also investing in the development of advanced 1500 F (816 C) capable noncontacting compliant finger seals to reduce leakage losses and have long life capability. A non-contacting 

brush/finger turbine seal (Ref. 11) designed by researchers at the NASA Glenn Research Center (GRC) is 

shown in Figure 26. Verification and refinement of design methodology using experimental data and 

analysis is on-going with testing up to 1472 F (800 C) and pressures up to 1.7 MPa differential 

(250 psid). The goal is to both validate design and analysis methodology, and to demonstrate low leakage 

and long life of the compliant non-contacting finger seals. This work is conducted in the GRC High 

Temperature, High Speed Turbine Seal Test Facility (Ref. 12), Figure 27. 

2.2 Unconventional Propulsion-Airframe Integration 

Several of the FW-funded system studies, most notably the MIT Double Bubble (Ref. 2) and NASA 

TurboElectric Distributed Propulsion (TeDP) (Ref. 5) aircraft concepts, Figure 28, departed from the 

conventional tube and wing configuration of current modern aircraft to capitalize on the potential for 

significant reduction in fuel/energy consumption afforded by propulors (i.e., fans) that are more highly 

integrated with the aircraft wing or fuselage to ingest the wing or fuselage boundary layer. According to 

the theory of Smith 9 (Ref. 13), propulsive efficiency, p, of an aircraft is increased when the propulsors 

are located such that they ingest the wing/fuselage boundary layer flow, Figure 29, where U0, Uj, and Uin

are the free stream velocity, jet exit velocity, and the inlet velocity, respectively. For the boundary layer 

ingesting (BLI) case, Uin  U0, the propulsive efficiency is higher compared to the non-ingesting case, Uin

= U0. The larger fraction of wing/fuselage boundary layer fluid ingested by the propulsors the greater the 

potential reduction in fuel/energy consumption. However, propulsors/fans will likely suffer a reduction in 

efficiency and stall margin when operating with a distorted inflow, which in this case will be continuous 

throughout the aircraft flight mission. Therefore, to maintain an overall system benefit the degradation in 

fan performance due to its interaction with a non-uniform inflow must be sufficiently less than the 

benefits obtained from ingesting the boundary layer to warrant the unconventional propulsion system/ 

airframe architecture. 
NASA/TM—2013-216548 15 

(a) (b) 

Figure 28.—(a) MIT Double Bubble aircraft concept with propulsors ingesting the fuselage boundary layer (Ref. 2), 

(b), NASA TurboElectric aircraft concept with propulsors ingesting the blended wing body boundary layer (Ref. 5). 

Figure 29.—Propulsive efficiency benefit from BLI versus conventional propulsion (Ref. 6). 

Based on industry studies which indicated potential for significant fuel burn reduction of BLI 

propulsion aircraft concepts (Refs. 14 to 16), NASA funded United Technologies Research Center 

(UTRC), via an NRA, to: (1) conduct system level studies of blended wing body aircraft with integrated 

BLI propulsion system architecture to assess potential fuel burn benefitsrelative to a baseline highperformance, pylon-mounted, propulsion system (Ref. 17), and (2) design a high-performance embedded, 

boundary layer ingesting, engine inlet and an associated distortion tolerant fan stage to minimize 

performance penalties of operating with continuous inlet distortion (Refs. 18 and 19). The results of the 

system studies indicated a potential 3 to 5 percent fuel burn reduction for a 5-engine configuration relative 

to an advanced baseline high bypass ratio turbofan engine, Figure 30, and also indicated that low-loss 

inlets and high-performance, distortion-tolerant turbomachines are key technologies to achieve the 

maximum potential fuel burn benefit. Subsequently, UTRC conducted a hierarchical, multi-objective, 

computational fluid dynamics-based aerodynamic design optimization that combined global and local 

shaping to design a high-performance embedded engine inlet (Ref. 18) and an associated distortion 

tolerant fan stage (Ref. 19) for testing in the NASA subsonic 8- by 6-Foot Supersonic Wind Tunnel 

(86 SWT). The UTRC CFD analysis of the optimized boundary layer ingesting inlet and distortion 

tolerant fan stage system predicts less than 2 percent reduction in fan efficiency relative to operation with 

uniform fan inflow conditions (Ref. 19). Wind tunnel experiments to validate the CFD prediction results 

are planned for 2015. The BLI inlet and distortion tolerant fan stage are shown installed in the NASA 86 

facility, including special test equipment consisting of a raised floor, boundary layer bleed system, and 

tunnel bump to control and vary the boundary layer profile entering the fan nacelle, and a NASA designed 

fast acting variable nozzle, Figure 31. NASA is also designing and fabricating rotatable multi-element 

rakes (circumferential and spanwise pressure and temperature probes) for the aerodynamic interface plane 

(AIP) and downstream of the exit of the fan exit guide vanes to measure fan performance parameters. An 

enlarged view of the distortion tolerant fan and layout of components to be installed in the NASA 86 

wind tunnel is provided in Figure 32. 

 
NASA/TM—2013-216548 16 

Figure 30.—Results from a system study indicating potential fuel burn benefits of BLL propulsion (Ref. 17). 

Figure 31.—Depiction of UTRC designed BLI inlet and distortion tolerant fan stage installed in 

NASA 86 wind tunnel which includes specialized test hardware. 

Figure 32.—Distortion tolerant fan stage layout of components for installation in the NASA 86 wind tunnel. 
NASA/TM—2013-216548 17 

(a) Inlet boundary conditions for 3D full-annulus unsteady CFD simulations. 

(b) Aero damping predicted from CFD simulations. 

Figure 33.—NASA high fidelity CFD for eromechanical analysis of UTRC designed 

distortion tolerant fan. 

NASA in-house researchers play a key role in being responsible for the high fidelity CFD analyses 

and assessments to determine that the fan is aeromechanically sound, and to identify critical speeds to 

avoid during testing. Such analysis requires extending the capability of their current CFD tools to enable 

aeromechanical analysis of a fan operating with continuous inlet distortion. This has entailed interpolating 

the fan inlet distortion from the UTRC simulation results to map their predicted distortion pattern to the 

NASA CFD code boundary conditions and then conducting unsteady full-annulus simulations to provide 

assessments of forced excitation and aeromechanical damping (Ref. 20), which included development of 

additional post-processing capability, Figure 33. Further planned improvements to the NASA CFD codes 

for aeromechanical analysis of boundary layer ingesting inlets coupled with distortion tolerant fan stages 

include incorporating CFD analysis of the boundary layer ingesting inlet with the distortion tolerant fan 

stage, and developing a closely coupled aerodynamic and structural analysis tool that would continuously 

update the structural analysis of the blade modes and deflections as the unsteady CFD simulations evolves 

and the fan blades move in and out of the distortion, rather than current practice where the structural 

analysis is first completed and the blade modes are fixed during the unsteady CFD simulations. 

 
NASA/TM—2013-216548 18 

Another topic of interest that the FW project is supporting research in developing adaptive fan blade 

technology, which involves training shape memory alloys for embedding in composite fan blades to effect 

optimal blade shape change to maximize fan efficiency at takeoff and cruise conditions, therefore 

minimizing fuel burn. Preliminary studies have shown that shape memory alloy wires will have more than 

adequate capability to achieve the required geometry changes when embedded in the composite. The use 

of shape memory alloys can also be applied to many other components, such as adaptive inlets, variable 

area nozzles, and active clearance control systems. The research also includes investigating carbon 

nanotube materials with the composite fan blades to toughen the blades. This will allow thinner fan blades 

that will improve aerodynamic efficiency and also be structurally tolerant to bird strikes. The adaptive fan 

blade research team is also developing analysis tools and methodology to validate the capability for static 

(non-rotating) testing of fan blades for impact resistance, alleviating the need for conducting multiple 

rotating fan blade tests of bird strike impact on the composite fan. 

2.3 Hybrid Gas-Electric Propulsion 

The N+3 system studies (Refs. 1 and 5) allude to the considerable promise of hybrid gas-electric 

propulsion to meet the aggressive goals for reducing fuel/energy consumption, emissions, and noise. 

Hybrid gas-electric propulsion includes concepts that are non-superconducting (e.g., Boeing SUGAR Volt 

(Ref. 1)) or superconducting (e.g., the NASA Turbo-Electric Distributed Propulsion concept, TeDP 

(Ref. 5)), and involve having both a gas turbine and electric motor (powered by batteries or other energy 

source) on the same shaft driving the fan/propulsors, or gas turbine engine generating power for electric 

motors driving the fan/propulsors. The FW-funded Boeing Subsonic Ultra Green Aircraft Research 

(SUGAR) studies (Ref. 1), conducted by Boeing in partnership with General Electric and Georgia Tech, 

showed the SUGAR Volt aircraft came the closest of all the concepts they studied to meeting the N+3 

goals with a 63 percent decrease in fuel burn and a 79 percent decrease in LTO NOx emissions. The 

SUGAR Volt involves a GE-developed idea for a hybrid electric-gas turbine engine (hFan) concept that 

uses batteries to assist during takeoff and cruise, development of a small ultra-high OPR core compressor, 

as well as other technology advancements, as shown in Figure 34. The additional power provided by the 

batteries could potentially allow sizing the gas turbine engine for cruise rather than takeoff. The gas 

turbine engine could then be powered down or perhaps even shut down during cruise with the battery 

powered electric motor driving the fans to propel the aircraft. There are no apparent show stoppers, but of 

course there are a number of technical challenges to address, including development of higher power 

density batteries and electric motors, before this concept is ready for commercial flight operations. 

The NASA Turbo-Electric Distributed Propulsion (TeDP) blended wing body aircraft concept also 

shows potential in meeting the NASA N+3 fuel/energy consumption, emissions, and noise goals, but 

requires considerable research and technology advancements to enable the cryogenic electric motors and 

components. The long-pole technology challenges requiring investments for the TeDP aircraft concept are 

illustrated in Figure 35. The FW project is currently focused on supporting the vehicle systems analysis of 

the TeDP (Refs. 21 and 22), and the development of the cryo-electric motors both through in-house research 

and an NRA with Advanced Magnetics Laboratory (AML). Other technologies indicated in Figure 35 are 

being pursued using NASA SBIRs or by leveraging investments made by the U.S. Air Force and Navy. 

The NRA with AML is developing high-fidelity modeling capability to design a low alternating 

current (AC) loss stator, and in-house efforts are aimed at developing processing of magnesium di-boride 

(MgB2) powders to fabricate 10 m diameter MgB2 superconducting filaments for motor windings. The 

FW project also plans in FY14 to augment funding for hybrid gas-electric propulsion technologies to 

include development of non-cyrogenic electric motors, initially aimed at doubling the current SOA in 

power density by 2020 time frame. 
NASA/TM—2013-216548 19 

Figure 34.—Overview of the GE “hFan” gas turbine-electric hybrid engine for the Boeing SUGAR Volt. 

Figure 35.—NASA turboelectric propulsion aircraft concept, and current technology investments. 

Another research topic of interest is aimed at investigating the challenges of managing and 

distributing potentially megawatt-level electric power for commercial aviation applications. To this end, 

the FW project recently funded a contract with Rolls-Royce North America to study the power grid 

architecture (Ref. 23) as well stability, transient response, control, and safety of a high-power electric grid 

(Ref. 24) for turboelectric propulsion aircraft. Rolls-Royce characterized and evaluated the critical 

dynamic and safety issues for the propulsion electric grid of a distributed propulsion system, and 

recommended system architecture and/or controls solutions that promote electrical stability, electric grid 

safety, and aircraft safety, and also assessed energy storage requirements. NASA in-house researchers at 

Glenn and Dryden are also supporting research to investigate power management and distribution 

challenges and issues for hybrid gas-electric propulsion concepts, including building scale bench-top or 

small test stands to understand the challenges of operating an electric motor and gas-turbine on the same 

shaft to power a fan. The promise of hybrid gas-electric concepts such as the Boeing SUGAR Volt and 

NASA TeDP for significant reductions in fuel/energy consumption, emissions, and noise makes this an 

attractive area for potential greater increase in NASA research investment in the future. 
NASA/TM—2013-216548 20 

3.0 Summary 

The FW Project is making investments in propulsion research and technology development that are 

expected to contribute to enabling significant reduction in fuel/energy consumption of future advanced fixed 

wing aircraft for the 2030 to 2035 time frame relative to current state-of-the-art engines. The investments in 

propulsion and materials research presented herein are to enable: (1) smaller higher OPR engine cores for 

increased thermal efficiency and to increase bypass ratio for increased propulsive efficiency, (2) robust, 

distortion tolerant fans for vehicle integrated boundary layer ingesting propulsion systems, and (3) 

cryocooled high power density electric motors for the TurboElectric Distributed Propulsion aircraft concept. 

Additional investments in hybrid gas-electric propulsion technologies are expected in the near future, which 

will include research and technology development of non-cryocooled (ambient temperature) high power 

density electric motors and components. The hybrid gas-electric and turbo-electric engine concepts show 

promise for significant reduction in fuel/energy consumption towards achieving the emissions and noise 

goals of the FW project. In FY14, the project is planning to augment investments in hybrid gas-electric 

propulsion toward development of non-cryogenic electric motors and components. 

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