Joseph E. Grady, William J. Haller, Philip E. Poinsatte, Michael C. Halbig, and Sydney L. Schnulo

Glenn Research Center, Cleveland, Ohio

Mrityunjay Singh

Ohio Aerospace Institute, Brook Park, Ohio

Don Weir, Natalie Wali, and Michael Vinup

Honeywell Aerospace, Phoenix, Arizona

Michael G. Jones

Langley Research Center, Hampton, Virginia

Clark Patterson and Tom Santelle

Rapid Prototyping+Manufacturing (rp+m), Avon Lake, Ohio

Jeremy Mehl

Princeton University, Princeton, New Jersey

A Fully Nonmetallic Gas Turbine Engine

Enabled by Additive Manufacturing

Part I: System Analysis, Component Identifi cation, Additive 

Manufacturing, and Testing of Polymer Composites

NASA/TM—2015-218748

May 2015
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Joseph E. Grady, William J. Haller, Philip E. Poinsatte, Michael C. Halbig, and Sydney L. Schnulo

Glenn Research Center, Cleveland, Ohio

Mrityunjay Singh

Ohio Aerospace Institute, Brook Park, Ohio

Don Weir, Natalie Wali, and Michael Vinup

Honeywell Aerospace, Phoenix, Arizona

Michael G. Jones

Langley Research Center, Hampton, Virginia

Clark Patterson and Tom Santelle

Rapid Prototyping+Manufacturing (rp+m), Avon Lake, Ohio

Jeremy Mehl

Princeton University, Princeton, New Jersey

A Fully Nonmetallic Gas Turbine Engine

Enabled by Additive Manufacturing

Part I: System Analysis, Component Identifi cation, Additive 

Manufacturing, and Testing of Polymer Composites

NASA/TM—2015-218748

May 2015

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors would like to thank Drs. Mike Dudley and Koushik Datta from NASA Aeronautics Research Institute for their 

continuous support and encouragements. We would like to thank test facility staff at NASA Glenn Research Center and NASA 

Langley Research Center for their help in IGVs and acoustic liner testing. The authors would also like to recognize Jeff Berton, 

Scott Jones and Mike Tong from the propulsion systems analysis organization at NASA Glenn Research Center who provided key 

technical support during the benefi t assessment effort. Technical help from various engineers from Honeywell and rp+m during 

this work is also gratefully acknowledged.

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This report is available in electronic form at http://www.sti.nasa.gov/ and http://ntrs.nasa.gov/
NASA/TM—2015-218748 1 

A Fully Nonmetallic Gas Turbine Engine 

Enabled by Additive Manufacturing 

Part I: System Analysis, Component Identification, 

Additive Manufacturing, and Testing of 

Polymer Composites 

Joseph E. Grady, William J. Haller, Philip E. Poinsatte, Michael C. Halbig, 

and Sydney L. Schnulo 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Mrityunjay Singh 

Ohio Aerospace Institute 

Brook Park, Ohio 44142 

Don Weir, Natalie Wali, and Michael Vinup 

Honeywell Aerospace 

Phoenix, Arizona 85034 

Michael G. Jones 

National Aeronautics and Space Administration 

Langley Research Center 

Hampton, Virginia 23681 

Clark Patterson and Tom Santelle 

Rapid Prototyping+Manufacturing (rp+m) 

Avon Lake, Ohio 44012 

Jeremy Mehl 

Princeton University 

Princeton, New Jersey 08544 

Abstract 

The research and development activities reported in this publication were carried out under NASA 

Aeronautics Research Institute (NARI) funded project entitled “A Fully Nonmetallic Gas Turbine Engine 

Enabled by Additive Manufacturing.” The objective of the project was to conduct evaluation of emerging 

materials and manufacturing technologies that will enable fully nonmetallic gas turbine engines. The 

results of the activities are described in three part report. The first part of the report contains the data and 

analysis of engine system trade studies, which were carried out to estimate reduction in engine emissions 

and fuel burn enabled due to advanced materials and manufacturing processes. A number of key engine 

components were identified in which advanced materials and additive manufacturing processes would 

provide the most significant benefits to engine operation. The technical scope of activities included an 

assessment of the feasibility of using additive manufacturing technologies to fabricate gas turbine engine 

components from polymer and ceramic matrix composites, which were accomplished by fabricating 

prototype engine components and testing them in simulated engine operating conditions. The 

manufacturing process parameters were developed and optimized for polymer and ceramic composites 

(described in detail in the second and third part of the report). A number of prototype components 

(inlet guide vane (IGV), acoustic liners, engine access door) were additively manufactured using high 
NASA/TM—2015-218748 2 

temperature polymer materials. Ceramic matrix composite components included turbine nozzle 

components. In addition, IGVs and acoustic liners were tested in simulated engine conditions in test rigs. 

The test results are reported and discussed in detail. 

Introduction 

Advanced lightweight and high temperature materials and technologies have made tremendous 

impact on the aerospace components and systems in terms of weight reduction, high operating 

temperature, reduced fuel burn, and reduced emissions. In particular, advances in composites materials 

and manufacturing technologies enabled the first polymer matrix composite (PMC) fan blades and fan 

containment system in gas turbine engines in the 1990s. More recent advances in ceramics and coatings 

will soon enable the first ceramic matrix composite (CMC) components in commercial aircraft engines in 

2016. Recent high TRL demonstration of CMC exhaust nozzle components will lead to introduction of 

CMC exhaust nozzles in commercial engines in the near future. Even with the introduction of PMCs and 

CMCs in gas turbine engines, the percentage of composites in these engines is relatively small, on the 

order of 10	to	20 percent, compared to nearly 50 percent for airframe structures, with most of it consisting 

of metallic components. 

As the temperature capability of PMCs increase from the state of the art (SOA) of 150 C (300 F) to 

350 C (660 F), the PMCs can be applied for a wide range of gas turbine structures. Similarly, the CMC 

technology continues to advance (Ref. 1). Advanced composite materials have created new opportunities 

for innovative component designs and new gas turbine designs/architectures that can take full advantage 

of the properties of the unique properties of composites. A wide variety of metallic components in the hot 

sections of current gas turbine engines could be replaced with ceramic matrix composites, reducing 

engine weight and the need for cooling air, and therefore increasing engine efficiency and reducing 

emissions and fuel burn. 

The goal of the current project, funded by NASA Aeronautics Research Institute (NARI), was to 

perform comprehensive evaluation of emerging materials and manufacturing technologies that will enable 

fully nonmetallic gas turbine engines. The technical activities were carried out to assess the feasibility of 

additive manufacturing technologies for fabricating complex polymer matrix composite (PMC) and 

ceramic matrix composite (CMC) gas turbine engine components. System analysis studies were 

conducted to assess the benefits of fully nonmetallic gas turbine engine in terms of fuel burn, emissions, 

reduction of part count, and cost. The potential benefits of this effort include: 

 Significant reduction in number of heavy metallic components, with more than 50 percent reduction 

of weight and the associated reduction of fuel burn/carbon dioxide emissions 

 Reduction in manufacturing costs by the simplified additive manufacturing of composite 

 Structural components, replacing casting of metals 

 Engine weight and part count reduction by reducing the need for joining and attachment of components 

 New engine architectures enabled by the unique properties of composites 

However, comprehensive engine design studies needs to be conducted to develop design and architecture 

concepts for a fully nonmetallic gas turbine engine that can be enabled by effective use of PMCs, CMCs 

and ceramic materials and integration of components through additive manufacturing of complex 

components. 

In order to demonstrate this concept, the effort was focused on 7000 lbf gas turbine engine. However, 

the materials and manufacturing technologies developed in this program are equally applicable to large 

gas turbine engine. The technical work was carried out by multidisciplinary, multiorganization 

NASA—industry team that included experts in engine design and analysis, and system analysis, additive 

manufacturing, polymers and PMCs, structural engineering, ceramic materials and CMCs, and 

sub-component testing under simulated engine conditions. 
NASA/TM—2015-218748 3 

System Analysis and Potential Benefits 

The systems analysis effort was tasked with evaluating the potential fuel burn, landing-takeoff (LTO) 

NOX and acoustic benefits of polymer matrix and ceramic matrix composite materials in a regional jet 

class engine. The first step in the process was the development of the current technology, small thrust 

(i.e., regional jet class) engine to serve as the baseline. After investigating the potential engine options and 

interaction with Honeywell, the initial decision was for the NASA systems analysis team to use an 

“AE3007-like” engine. The process began with a review of the existing Numerical Propulsion System 

Simulation (NPSS) engine model (Refs. 2 to 4) and several issues were discovered. The model was 

constructed several years ago and needed significant updating to work with the current version of NPSS. 

The team then decided that an alternate engine cycle was needed. The systems analysis organization had a 

functioning NPSS model of a CF34-8C “like” engine, along with an airplane model of the CRJ-900 

regional jet that uses the CF34 engine. Both models were created by Georgia Tech during the 

development of the FAA-sponsored Environmental Design Space (EDS capability) (Ref. 5). A key part of 

the EDS development process involved industry feedback on the model assumptions. Although the 

CF34-8C is a larger thrust engine than the AE3007, it does reside in the regional jet thrust class. After 

consultation with the other team members, and with Honeywell’s concurrence, the decision was made to 

proceed with the CF34-8C as the baseline. Figure 1 shows the baseline engine and airplane data used for 

the system analysis of engine (CF34-8C5 “like”) and the airframe (CRJ900LR “like”). 

Figure 1.—Key baseline engine and airplane data with design mission for current technology engine 

(CF34-8C5 “like”) and the airframe (CRJ900LR “like”). 

 
NASA/TM—2015-218748 4 

Figure 2.—Regional jet fuel burn sensitivities showing fuel burn reduction and 

engine weight reduction. 

The analysis effort commenced with the generation of initial performance and weight benefits from 

inserting polymer matrix composites (PMC) and ceramic matrix composites (CMC) into the baseline 

engine. Based on input from the materials experts, it was decided that the maximum allowable material 

temperature for the PMCs was 415 °F). As such, PMCs were envisioned for use in several locations in the 

front portion of the engine (inlet acoustic duct, fan stator, and the first four rows of high-pressure 

compressor vanes). In the engine’s hot section, 2600 °F CMCs were inserted in the combustor liner, highpressure turbine (HPT) vanes/blades, low-pressure turbine (LPT) vanes/blades and core nozzle. The cycle 

was reoptimized to eliminate all turbine blade/vane cooling. The “advanced” cycle produced ~2.6 percent 

improvement in specific fuel consumption. The inclusion of the composite materials enabled the engine 

weight to decrease by ~14.5 percent. Figure 2 shows the analysis of regional jet fuel burn sensitivities 

indicating influence of engine weight reduction on the reduction in fuel burn. The combination of fuel 

efficiency improvement and engine weight reduction would indicate an airplane that has a 4.9 percent 

reduction in total fuel burn versus the current technology baseline.

The LTO NOX benefit was derived primarily from the drop in combustor exit temperature with the 

addition of the uncooled, CMC turbine vanes and blades. The baseline engine had a margin to the 

CAEP/6 stringency of 20.4 percent. The “advanced” engine has a margin of 28.7 percent, an 8.3 percent 

increase. There were no direct acoustic benefits identified for the PMC and CMC materials. There was no 

change in the fan tip speed and minimal differences in the jet velocity versus the baseline engine. As such, 

it was determined that the “advanced” engine/airplane noise would be almost identical to the baseline. 

However, during the evaluation a potential noise reduction opportunity was discussed. Taking advantage 

of additive manufacturing could enable the production of variable depth inlet acoustic liners. This tailored 

approach could capture the larger tones and discrete frequencies and as such would provide an acoustic 

benefit. NASA acoustic discipline experts, in consultation with Honeywell, plan to work with the systems 

analysis team to develop a benefit from such a design. If the benefits can be quantified, a full system 

acoustic assessment could be conducted. 

The second part of the assessment was to allocate the projected 4.9 percent fuel burn benefit between 

the PMCs and CMCs. Table 1 shows the influence of different materials and components on the system 

level benefits. In addition to the weight reduction versus the metallic parts replaced, CMC usage enabled 

the elimination of all turbine blade and vane cooling. By removing the need for cooling and the associated 

mixing losses, an efficiency improvement in the HPT (1.5 pts) and LPT (0.5 pt) was assumed. The PMC 

benefits were entirely weight reduction based and had no impact on the engine performance. As such, the 

vast majority (~90 percent) of the fuel burn improvement is attributed to the hot section CMCs. 

 
NASA/TM—2015-218748 5 

TABLE 1.—EFFECT OF KEY MATERIALS TECHNOLOGIES ON VARIOUS SYSTEM LEVEL BENEFITS 

In summary, an assessment was performed by NASA’s propulsion systems analysis branch to 

quantify the potential benefits of PMC and CMC engine components that could be produced via additive 

manufacturing processes for a regional-jet class system. The study assumed the availability of 2600 °F 

CMCs and 415 °F PMCs. The CMCs were applied in the combustor liner, the high-pressure and 

low-pressure turbine blade/vanes and the core nozzle. The benefits included a weight reduction in the 

aforementioned components due to replacing metallic parts with ceramic-based alternatives, in addition to 

the elimination of requisite turbine cooling. The PMCs were applied in the inlet acoustic liner, the fan 

stator and the first four rows of the high-pressure compressor vanes. The PMC benefits consisted of 

weight reductions only in each component. The resultant propulsion system generated a 4.9 percent fuel 

burn improvement over the baseline and a 7.7 percent increase in landing-takeoff (LTO NOX) margin 

with respect to the CAEP/6 stringency. The assessment estimated minimal acoustic impact. Finally, there 

may be some durability concerns with the use of composites in the turbine rotors and the fan stator that 

could merit further investigation. 

Identification of Potential Components 

Polymer Matrix Composites 

During the initial phases of the project, several target components for the PMC manufacturing efforts 

were identified. These selections were made based on the operational requirements, material properties, 

manufacturing capability of commercially available and developmental polymer composites. The initial 

component list was narrowed to two types of components for the Phase I demonstration. Figure 3 shows few 

of the high-payoff engine components identified for advanced materials and manufacturing technologies 

Polymer based components included inlet guide vanes (IGVs) and acoustic liners. The inlet guide vane 

(IGV) is a static structural component currently made of Ti-6Al-4V alloy. This component is subjected to 

low pressures and temperatures suitable for the polymer matrix composite material systems assessed in this 

program. There has been industrial experience with manufacturing and testing a PEEK thermoplastic IGV. 

The polymer composite IGV components are expected to provide benefits from effective weight reduction, 

superior strength and temperature capability as well as advanced manufacturing processes of this program. 
NASA/TM—2015-218748 6 

Figure 3.—High-payoff engine components identified for advanced 

materials and manufacturing technologies. 

Acoustic attenuation materials are used in almost all components of a turbofan engine. Acoustic liner 

operating temperature is 250 °F for a typical propulsion engine aft-fan duct. Additive manufacturing may 

be used to make absorptive structures not feasible with conventional techniques. Previous efforts have 

been focused on locally-reacting resonator liners. This project will help to determine if it is possible to 

build advanced liners using passive-destructive interference or direct dissipation of the acoustic waves. 

Traditional manufacturing methods typically used to fabricate locally-reacting resonator acoustic 

liners include fabrication of honeycomb resonator structures using sheet metal forming methods, 

fabrication of a porous facesheet using ultrasonic or laser drilling methods to form perforations in a sheet 

metal cover, and bonding facesheet to the honeycomb or bulk absorber using well known metal brazing or 

welding methods. Additive manufacturing processes combine all of these steps into a single process in 

which the component is fabricated by progressively building the structural element layer by layer, in a 

process driven by a geometric file describing the dimensions of the structure. Acoustic liner fabrication 

using traditional and an additive manufacturing approach is shown in Figure 4. No joining is needed and 

the required porosity is “built-in” to the facesheet without any need for post-fabrication drilling or 

machining. Novel acoustic liner concepts may also be constructed. This process has the added potential 

for reduced weight and fabrication cost. 

The technical approach for evaluation of PMC acoustic liners fabricated by Fused Deposition 

Modeling included: 

 Selection of a conventional liner sample from a previous NASA program (Ref. 6). 

 Fabrication of a conventional locally reacting liner sample using additive manufacturing. 

 Build and test the advanced liner concept shown in Figure 5, which is not feasible with conventional 

manufacturing techniques, using Fused Deposition Modeling. 

 Test three samples in the NASA Langley Grazing Flow Impedance Tube (GFIT) shown in Figure 5. 

 Fabricate a typical acoustically-treated access panel for a turbofan engine bypass duct. 
NASA/TM—2015-218748 7 

Figure 4.—Additive manufacturing simplifies acoustic liner fabrication. 

Figure 5.—Fused deposition modeling enables fabrication of advanced passive-destructive 

acoustic liner concept. 

Ceramic Matrix Composites 

A number of components were identified for the Phase I ceramic manufacturing effort. A balance had 

to be struck between identifying the components that would offer nontrivial fuel burn reduction if 

manufactured as ceramics and identifying the components that were relatively low-stress. Other factors 

included manufacturing complexity (e.g., geometry and cooling passage schemes), part size and 

structures/attachments. 

The high pressure turbine second stage nozzle doublet vane was chosen because it is a static structural 

component subjected to high temperatures yet it is in a relatively low stress environment. The current 

state of the art for the nozzle is a multipiece alloy assembly. Individual doublets need to be stressrelieved, brazed, aged and inspected during the manufacturing process, then tack welded and brazed again 

into the assembly. A benefit from this program would be to reduce manufacturing time and labor. In 

addition, cooling air to the nozzle may be reduced. 
NASA/TM—2015-218748 8 

A second component is the high pressure turbine shroud for auxiliary power units (APUs). In this 

unit, the shroud sees a maximum gas temperature of 1800 °F. The shroud is the most ideal static 

component as it offers great potential reductions in SFC (e.g., blade tip clearance and weight reduction) 

life cost, and blade excitation over the current segmented metallic shroud. There has been significant 

confidence in the design and test of ceramic shrouds with several successful programs, including the 

Small Heavy Fuel Engine (SHFE) program at Honeywell. 

Additive Manufacturing of Polymer Components 

In	this	project,	Fused Deposition Modeling (FDM) was used as additive manufacturing technology 

for fabrication of polymer composites. Different types of materials were used including commercially 

available polyetherimides—Ultem 9085 and experimental Ultem 1000 mixed with 10 percent chopped 

carbon fiber. Detailed experimental results of the polymer composite process development and 

characterization has been described in the second report. In this report, we will describe the testing details 

of inlet guide vanes fabricated using ABS and carbon fiber reinforced Ultem 1000 and acoustic liners 

fabricated from Ultem 9085.

Designs for two alternate liner geometries were developed for manufacturing trials. A Helmholtz 

design similar to the current baseline honeycomb concept and the advanced “Passive-Destructive” liner 

concept, developed at Honeywell and based on Herschel-Quincke tube theory, were both fabricated for 

acoustic evaluation. Initial design iterations for both designs were fabricated by Fused Deposition 

Modeling. FDM process parameters were optimized to print the thinnest acceptable septum walls and best 

geometric cell shapes for acoustic attenuation and light weight. Two Helmholtz samples were delivered to 

NASA for acoustic testing. These samples were printed in Ultem 9085, which was chosen as the best 

alternative material based on the anticipated 250 °F operating temperature of the acoustic liner in engine 

operation. Figure 6 shows the pictures of the liners and Grazing Flow Impedance Tube test rig at NASA 

Langley Research Center (LaRC). Details of the test results have been discussed in following sections. 

Figure 6.—Noise attenuation of baseline and advanced acoustic liners was evaluated in 

NASA LaRC grazing flow impedance tube acoustic facility. 
NASA/TM—2015-218748 9 

Testing of Additively Manufactured Polymer Components 

Inlet Guide Vane Cascade Testing 

A set of additively manufactured inlet guide vanes were tested in the Engine Research Building 

(ERB) SW-2 wind tunnel facility as shown in Figures 7(a) and (b). This facility uses the lab-wide central 

exhaust system to draw atmospheric air through an acrylic wind tunnel. The test section consists of a 

linear cascade of nine inlet guide vanes spaced 1.5 in. apart. The test area cross-section was 4.5 in. high 

by 15 in. wide. Flow rate was controlled with a throttling valve downstream of the test section. Inlet flow 

conditioning tunnel sections were removed for this experiment to allow unobstructed camera views of the 

test section. 

(a) 

(b) 

Figure 7.—(a) NASA ERB SW-2 general wind tunnel (Schematic). (b) ERB SW-2 

facility with IGV cascade and DIC system. 
NASA/TM—2015-218748 10 

The four different inlet guide vane designs that were tested are shown in Figure 8. Two vanes were 

made of ABS and two were made of Ultem 1000 with 10 percent chopped carbon fiber. Additionally two 

fillet designs at the hub and tip of each vane were employed in this test. One set of vanes (ABS and 

carbon fiber Ultem) was fabricated with the standard Honeywell fillets, while another set of vanes 

(ABS and Carbon fiber Ultem) was fabricated with a larger reinforced fillet at the tip and hub. The vanes 

are roughly 6 in. long, with an airfoil section 4.5 in. long. The chord varies from 1.25 to 2 in. All vanes 

were printed with buildup proceeding from the leading edge to the trailing edge (right to left in the vane 

configuration shown in Figure 9). Each vane was later painted with a speckled pattern to allow for 

imaging by the ARAMIS DIC system described below. 

 

 

Figure 8.—Inlet guide vanes tested in SW2 cascade. 

Figure 9.—Inlet guide vane cascade test section. 
NASA/TM—2015-218748 11 

The individual vanes were mounted in the tunnel ceiling and floor and were secured with a screw 

through the tunnel ceiling and threaded into the vane stem. The vanes could be rotated and fixed at 

different angles of incidence. The tunnel inlet was nominally 4.5 by 15 in. The test section included nine 

vanes total with vanes 4 to 7 being the test articles. The inlet airstream was at room temperature and 

passed through no flow conditioning sections upstream of the test cascade. 

Strain under an aerodynamic load was measured on several vanes using an optical stereoscopic 

Digital Image Correlation (DIC) technique. All the test vanes were painted with a speckled pattern. The 

two DIC cameras were placed looking down stream into the tunnel inlet. They were arranged in such a 

way as to image the pressure face of the four center vanes. The four different test designs were located in 

this view. Comparison of the speckled patterns of two camera views of the same vane yields the 

deformation of each vane as it bends under changing loads. 

Tests were performed at various inlet velocities ranging from nominally 100 to 600 ft/sec and at 

angles from 0 to 60. Additionally the vane cascade was tested in a “closed” position i.e., vanes were set 

to 90 to the inlet flow in an attempt to fracture the vanes. Deflection up to several millimeters was 

observed but no fracture was detected, even at the “closed” condition. Deflection and strain data from the 

DIC system is summarized in Figure 10. 

Figure 10 shows typical displacement data acquired from the ARAMIS DIC system. It shows the 

displacement in the Z (normal to pressure face) direction of the four test vanes at 20, 30, 40, and 50

incidence angle under a 400 ft/sec an airstream. The vanes numbered from left to right are listed in the 

Table 2. As expected, for all vanes the displacement increases with increasing incidence angle. Also, the 

largest displacement generally occurs on the ABS vane with the baseline fillet design (vane 5). The vane 

maximum displacement data for nominal velocities from 100 to 600 ft/sec and 20, 30, 40 and 50

angles are summarized in Figure 11. From the graph it is clearly evident that for each angle, the maximum 

deflection increases with increasing velocity. Also evident, as previously mentioned, the max deflection 

Figure 10.—Displacement in the Z-direction for 20, 30, 40 and 50 incidence angle at 

400 ft/sec velocity. 
NASA/TM—2015-218748 12 

TABLE 2.—POSITION OF TEST VANES IN CASCADE 

Vane number 

(left to right) 

Material Fillet design 

4 ABS Reinforced (RP+M) 

5 ABS Baseline (Honeywell) 

6 Ultem 1000 w/Carbon Baseline (Honeywell) 

7 Ultem 1000 w/ Carbon Reinforced (RP+M) 

Figure 11.—Displacement data at different velocities and incident angle showing maximum displacement in 

Z-direction. 

increases with increasing incidence angle. Generally the ABS vanes (4 and 5) show higher deflection than 

the carbon reinforced Ultem 1000 vanes (6 and 7). Additionally this data seems to show that that 

reinforced fillet improved the stiffness of the vane for the ABS cases but not so much for the Ultem cases. 

It is important to note however, that these tests were conducted with a very limited number of trials and 

test specimens and more samples should be tested to confirm this result. 

Similarly, Figures 12 and 13 illustrate typical strain data acquired in this test. Figure 12 represents 

strain in the span wise direction for the four test vanes at 20, 30, 40, and 50 incidence angle under a 

400 ft/sec an airstream. Figure 13 shows maximum positive and negative strain in the span wise direction 

as a function of velocity for the 50 angle case. As with the displacement, the highest strain values were 

observed on vane 5 (ABS Baseline fillet). The strain values do exhibit a somewhat streaky pattern in all 

cases. Typically a large negative strain is seen near mid-span of blade and a large positive strain near hub 

especially at higher angles and speeds. Again, these tests were conducted with a very limited number of 

runs and test specimens. The main objective was to determine if the vanes would could withstand in air 

speeds that may be typical of an actual IGV. While a fair amount of bending was observed, none of the 

four center test vanes fractured during these tests. 
NASA/TM—2015-218748 13 

Figure 12.—Strain in the Y-direction for 20, 30, 40 and 50 incidence angle at 400 ft/sec velocity. 

Figure 13.—Maximum strain in Y-direction versus velocity at 50 angle. 

 
NASA/TM—2015-218748 14 

Rig Testing of Acoustic Liners 

Acoustic liner testing was conducted in the NASA Langley Research Center Grazing Flow 

Impedance Tube (GFIT) acoustic rig. The measurement process is documented in Reference 3. Acoustic 

data were collected at three values of flow Mach number and two tone amplitudes. Tones at frequencies 

of 0.4 to 3.0 kHz in increments of 0.2 kHz were measured. Table 3 provides a summary of the test matrix. 

The focus of the analysis was on the no flow and the Mach number of 0.3 conditions. Configuration F3 

was created at NASA Langley by adding a 3 cgs rayl wire mesh cover sheet to reduce generated flow 

noise due to the size of the surface openings. 

Typical measurement results are shown in Figure 14 for the conventional construction liner for a tone 

source level of 120 dB with no flow. Comparable data is available for all of the conditions summarized in 

Table 3. The graph shows the changes in tone amplitude from the upstream to the downstream end of the 

liner sample. The significant change in the tone amplitude occurs at frequencies of 1.4 and 1.6 kHz 

corresponding to the resonant frequency of this locally-reacting liner. 

A rough estimate of the insertion loss for each liner sample can be made by subtracting the sound 

level at 40 in. from the sound level at 0 in. Figure 15 shows the results of that calculation for the no flow 

and Mach number of 0.3 cases at 120 dB tone level. The results at 140 dB are similar. 

TABLE 3.—CONFIGURATIONS TESTED AT THE NASA LANGLEY GRAZING FLOW IMPEDANCE TEST FACILITY 

Configuration Liner Mach = 0.0 Mach = 0.3 Mach = 0.5 

120 dB 140 dB 120 dB 140 dB 120 dB 140 dB 

1 Conventional construction X X X X X X 

2 Fused deposition model X X X X X X 

3 Advanced concept X X X X X X 

F3 Advanced concept w/linear facesheet X X X X X -- 

Figure 14.—Measured sound levels of the conventional acoustic liner. 

 

40

50

60

70

80

90

100

110

120

130

140

0 5 10 15 20 25 30 35 40 45

SPL, dB

Axial Location, In

, ,

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

2.4

2.6

2.8

3.0

Frequency, kHz
NASA/TM—2015-218748 15 

Figure 15.—Insertion loss of all liner samples estimated from the GFIT measurements. 

For the no flow case (solid lines in Fig. 15), the resonance frequencies of the conventional 

construction (Configuration 1) and the Fused Deposition Model (Configuration 2) are comparable. The 

frequency of the Fused Deposition Model is slightly higher (of the order of 100 Hz), indicating that the 

average chamber height of the liner may be slightly shorter than the conventional construction liner. The 

advanced concept liner data (Configuration 3 and F3) do not show evidence of a resonant frequency as 

expected. The advanced concept liner insertion loss in the range of 1600 to 3000 Hz shows improved 

higher frequency attenuation per the design intent. 

The Mach 0.3 data (dotted lines in Fig. 15) show that the flow effect on the liner samples differ. The 

flow had minimal effect on the performance of the advanced concept liners. However, the flow effect had 

a significant effect on the performance of the conventional liner samples. The conventional construction 

liner lost its resonance behavior, but the Fused Deposition Model liner maintained a resonant behavior, 

but it shifted to a higher frequency. The flow also improved the attenuation of the conventional liners 

from 1800 to 2600 Hz, eliminating the benefit of the advance liner construction. 

The acoustic impedance of the four liner samples was educed following the process in Reference 3. 

Figure 16 shows the results for the two locally-reacting liner configurations. Also shown on the figure are 

the results from the DC flow measurements reported in Reference 2. As expected, the GFIT measured 

resistance of the conventional construction liner matched the DC Flow results reasonably well. However, 

the resistance of the Fused Deposition Model liner with flow was noticeably less than the conventional 

construction liner. This result explains the insertion loss results in Figure 15. Lower resistance due to flow 

results in lower high frequency attenuation and maintains the resonant characteristic of the liner. It is 

likely that the hole diameter of the Fused Deposition Model liner was larger than the conventional 

construction liner. Measurements are planned to verify the diameter of the holes in the two liners. 

 

0

10

20

30

40

50

400 800 1200 1600 2000 2400 2800 3200

Insertion Loss, dB

Frequency, Hz

SAP1M000L2

SAP1M300L2

SAP2M000L2

SAP2M300L2

SAP3M000L2

SAP3M300L2

SAPF3M000L2

SAPF3M300L2

Key: 

SAP1M000L2 = Configuration 1, Mach = 0.0 SAP3M000L2 = Configuration 3, Mach = 0.0 

SAP1M300L2 = Configuration 1, Mach = 0.3 SAP3M300L2 = Configuration 3, Mach = 0.3 

SAP2M000L2 = Configuration 2, Mach = 0.0 SAPF3M000L2 = Configuration F3, Mach = 0.0 

SAP2M300L2 = Configuration 2, Mach = 0.3 SAPF3M300L2 = Configuration F3, Mach = 0.3 
NASA/TM—2015-218748 16 

(a) 

(b) 

Figure 16.—Acoustic impedance of the conventional liner as educed from the GFIT measurements. 

(a) Conventional construction. (b) FDM liner. 

-6

-4

-2

0

2

4

400 800 1200 1600 2000 2400

Normalized Impedance

Frequency, Hz

R,M0,L2

R,M3,L2

R,M0,L4

R,M3,L4

X,M0,L2

X,M3,L2

X,M0,L4

X,M3,L4

DC Res M0

DC Res M3

-6

-4

-2

0

2

4

400 800 1200 1600 2000 2400

Normalized Impedance

Frequency, Hz

R,M0,L2

R,M3,L2

R,M0,L4

R,M3,L4

X,M0,L2

X,M3,L2

X,M0,L4

X,M3,L4

DC Res M0

DC Res M3

Key: 

R,M0,L2 = Resistance, Mach = 0.0, SPL = 120 dB X,M0,L2 = Reactance, Mach = 0.0, SPL = 120 dB 

R,M3,L2 = Resistance, Mach = 0.3, SPL = 120 dB X,M3,L2 = Reactance, Mach = 0.3, SPL = 120 dB 

R,M0,L4 = Resistance, Mach = 0.0, SPL = 140 dB X,M0,L4 = Reactance, Mach = 0.0, SPL = 140 dB 

R,M3,L4 = Resistance, Mach = 0.3, SPL = 140 dB X,M3,L4 = Reactance, Mach = 0.3, SPL = 140 dB 

DC Res M0 = Resistance at Mach = 0.0, Ref. 7 DC Res M3 = Resistance at Mach = 0.3, Ref XX1 
NASA/TM—2015-218748 17 

Figure 17 shows the results for the two passive-destructive liner concepts. First, it is noted that the 

liner resistance is higher than for the locally reacting configurations with no flow. This higher resistance 

leads to the improved broadband performance at the no flow condition. Further, it can be seen that the 

advanced liner shows minimal change in resistance with flow. Since the resistance of the advanced liner 

is comparable to the resistance of the conventional liner at Mach 0.3, it is expected that acoustic 

performance of the two liners will be similar. Further optimization of the advance liner concept is needed 

to exploit the characteristics that were identified in this study. 

(a) 

(b) 

Figure 17.—Acoustic impedance of the advanced concept liner as educed from the GFIT measurements. 

(a) As fabricated. (b) With linear facesheet. 

-6

-4

-2

0

2

4

400 800 1200 1600 2000 2400

Normalized Impedance

Frequency, Hz

R,M0,L2

R,M3,L2

R,M0,L4

R,M3,L4

X,M0,L2

X,M3,L2

X,M0,L4

X,M3,L4

-6

-4

-2

0

2

4

400 800 1200 1600 2000 2400

Normalized Impedance

Frequency, Hz

R,M0,L2

R,M3,L2

R,M3,L4

X,M0,L2

X,M3,L2

X,M3,L4

Key: 

R,M0,L2 = Resistance, Mach = 0.0, SPL = 120 dB X,M0,L2 = Reactance, Mach = 0.0, SPL = 120 dB 

R,M3,L2 = Resistance, Mach = 0.3, SPL = 120 dB X,M3,L2 = Reactance, Mach = 0.3, SPL = 120 dB 

R,M0,L4 = Resistance, Mach = 0.0, SPL = 140 dB X,M0,L4 = Reactance, Mach = 0.0, SPL = 140 dB 

R,M3,L4 = Resistance, Mach = 0.3, SPL = 140 dB X,M3,L4 = Reactance, Mach = 0.3, SPL = 140 dB 
NASA/TM—2015-218748 18 

Summary and Conclusions 

System analysis studies indicate that there are potential fuel saving benefits from PMC and CMC 

engine components for a regional-jet class system. The CMC components were combustor liner, the 

high-pressure and low-pressure turbine blade/vanes, and the core nozzle. The benefits included a weight 

reduction in the aforementioned components due to replacing metallic parts with ceramic-based 

alternatives, in addition to the elimination of requisite turbine cooling. The PMCs were applied in the 

inlet acoustic liner, the fan stator and the first four rows of the high-pressure compressor vanes and the 

PMC benefits consisted of weight reductions only in each component. The resultant propulsion system 

generated a 4.9 percent fuel burn improvement over the baseline and a 7.7 percent increase in landingtakeoff (LTO NOX) margin with respect to the CAEP/6 stringency. 

In case of polymer composites, two types of acoustic liners were fabricated from Ultem 9085 polymer 

and tested in rigs. They included a Helmholtz design similar to the current baseline honeycomb concept 

and the advanced “Passive-Destructive” liner concept, developed at Honeywell and based on HerschelQuincke tube theory. Compressor inlet guide vanes, fabricated from ABS and carbon fiber reinforced 

Ultem 1000, were tested in wind tunnel to measure the deflection and strain. The ABS vanes show higher 

deflection than the carbon reinforced Ultem 1000 vanes. Additionally this data seems to show that 

reinforced fillet improved the stiffness of the vane for the ABS but not so much for the Ultem 1000. 

Detailed analysis of results for polymer and ceramic composites are presented in second and third part of 

the report. 

References

1. M.C. Halbig, M.H. Jaskowiak, J.D. Kiser, and D. Zhu, “Evaluation of Ceramic Matrix Composite 

Technology for Aircraft Turbine Engine Applications,” 51st AIAA Aerospace Sciences Meeting 

including the New Horizons Forum and Aerospace Exposition, 07–10 January 2013, Grapevine 

(Dallas/Ft. Worth Region), Texas. 

2. R.W. Claus, A.L. Evans, J.K. Lytle and L.D. Nichols, “Numerical Propulsion System Simulation,” 

Computing Systems in Engineering, Vol. 2, No. 4, pp.357–364, 1991. 

3. NPSS User Guide Software Release: NPSS_1.6.5. 

4. NPSS Reference Sheets Software Release: NPSS_1.6.5. 

5. M.R. Kirby and D.N. Mavris, “The Environmental Design Space,” 26th Congress of International 

Council of the Aeronautical Sciences (ICAS), Anchorage, Alaska, Sept 14–19, 2008, ICAS 2008-

4.7.3. 

6. Asif A. Syed, et al, “The Steady Flow Resistance of Perforated Sheet Materials in High Speed 

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7. M.G. Jones, et al., “Comparative Study of Impedance Education Methods, Part 2: NASA Tests and 

Methodology.” AIAA-2013-2125, 19th AIAA/CEAS Aeroacoustics Conference, May 27–29, 2013, 

Berlin, Germany.