Russell W. Claus

Glenn Research Center, Cleveland, Ohio

Tim Beach

Coyote Hollow Consulting, LLC, Walton Hills, Ohio

Mark Turner and Kiran Siddappaji

University of Cincinnati, Cincinnati, Ohio 

Eric S. Hendricks

Glenn Research Center, Cleveland, Ohio

Geometry and Simulation Results for a Gas Turbine 

ReSresentatiYe of tKe (nerJy (f¿Fient (nJine (((

NASA/TM—2015-218408

February 2015
NASA ST, 3roJram    in 3ro¿le

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Russell W. Claus

Glenn Research Center, Cleveland, Ohio

Tim Beach

Coyote Hollow Consulting, LLC, Walton Hills, Ohio

Mark Turner and Kiran Siddappaji

University of Cincinnati, Cincinnati, Ohio 

Eric S. Hendricks

Glenn Research Center, Cleveland, Ohio

Geometry and Simulation Results for a Gas Turbine 

ReSresentatiYe of tKe (nergy (f¿Fient (ngine (((

NASA/TM—2015-218408

February 2015

National AeronautiFs and

SSaFe Administration

Glenn ResearFK &enter 

&leYeland 2Kio 4415
Acknowledgments

TKis researFK Zas suSSorted by tKe Fi[ed-:ing element of tKe Fundamental AeronautiFs Program at NASA TKe autKors Zould 

liNe to furtKer aFNnoZledge tKe assistanFe of &Kris Miller %ill +aller Greg Follen and SFott ToZnsend at NASA Glenn TKis 

reSort Zould not be Sossible ZitKout tKe suSSort of AndreZ %ree]e-StringfelloZ and Pat NisNode retired at General (leFtriF

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AYailable from

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Geometry and Simulation Results for a Gas Turbine 

Representative of the Energy Efficient Engine (EEE) 

Russell W. Claus 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Tim Beach 

Coyote Hollow Consulting, LLC 

Walton Hills, Ohio 44146 

Mark Turner and Kiran Siddappaji 

University of Cincinnati 

Cincinnati, Ohio 45220 

Eric S. Hendricks 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This paper describes the geometry and simulation results of a gas-turbine engine based on the original 

EEE engine developed in the 1980s. While the EEE engine was never in production, the technology 

developed during the program underpins many of the current generation of gas turbine engines. This 

geometry is being explored as a potential multi-stage turbomachinery test case that may be used to 

develop technology for virtual full-engine simulation. Simulation results were used to test the validity of 

each component geometry representation. Results are compared to a zero-dimensional engine model 

developed from experimental data. The geometry is captured in a series of Initial Graphical Exchange 

Specification (IGES) files and is available on a supplemental DVD to this report. 

Nomenclature 

BPR By Pass Ratio 

CAD Computer Aided Design 

EEE Energy Efficient Engine 

HPC High Pressure Compressor 

HPT High Pressure Turbine 

ICLS Integrated Core Low Spool 

IGES Initial Graphical Exchange Specification 

LPC Low Pressure Compressor 

LPT Low Pressure Turbine 

Nc Corrected rotational speed 

NLH Non-Linear Harmonic 

NPSS Numerical Propulsion System Simulator 

1.0 Introduction 

NPSS (Numerical Propulsion System Simulator) was a NASA program created in the early 1990s to 

build a virtual engine simulation capability (Ref. 1). Leveraging advanced computing technologies, its 

NASA/TM—2015-218408 1 
goal was to develop a framework that enabled reduced development timeframes based on virtual engine 

simulation. The project developed research efforts that explored multi-fidelity, multidiscipline models of 

engine systems to baseline computational accuracies and address fundamental questions exploring the 

practicality of the technique (Ref. 2). Today, NPSS is used as a zero-dimensional engine model 

simulation environment supporting the development of many new propulsion systems. 

Beyond a few demonstrations (Refs. 2, 3, and 4) of full engine simulations, the practical application 

of NPSS remains as a zero-dimensional cycle model. Until recently, a full, three-dimensional simulation 

of a gas turbine engine was not a practical capability for any engineering organization. Today, many 

advances make full-engine simulations practical. First, computational resources for these simulations may 

be demanding but the advance of chip and software technology has greatly mitigated this challenge. 

Second, many engine component simulations employ proprietary codes and models which limit the utility 

of the simulation. Commercial codes, while not fully comprehensive, are making significant progress to 

reduce this limitation. A third limitation is the availability of geometry to test and improve component 

codes in a full system environment. Previous studies have used proprietary geometries that could not be 

further disseminated. 

This report publishes geometry files from a series of reports (Refs. 5, 6, and 7) and the original 

Compter Aided Design (CAD) derived drawings. The geometry is captured in a series of Initial Graphical 

Exchange Specification (IGES) files. Some of these IGES files trace to the original CAD drawings for the 

EEE turbomachinery. Some of the files (specifically for the HPC) were not created from the original 

CAD drawings, but were recreated from the data in the NASA contractor report (Ref. 5). The original 

CAD-based files were used in a series of simulations by Hall (Ref. 4, 8, and 9). 

A series of three-dimensional numerical simulations were conducted to validate the geometry. The 

High Pressure Compressor (HPC) was the most challenging component to simulate and the geometry 

from the two different sources provided significantly different results. A recommendation for component 

simulations is provided based on the calculations for each turbomachinery component of the engine. It is 

shown that these results, which should be considered preliminary, may be adequate to assemble into a full 

gas-turbine simulation. Additionally, it is hoped that the geometries can be used as multi-stage benchmark 

testcases to further the development of new models and simulation techniques. 

2.0 EEE Geometry 

In the 1980s, NASA funded the Energy Efficient Engine program to demonstrate fuel efficient 

designs for the next generation of transport aircraft. New low-emission combustor designs were tested to 

meet emissions goals. For the turbomachinery, the primary focus was to meet a 12 percent reduction in 

installed specific fuel consumption. Both Pratt & Whitney and General Electric (GE) were awarded 

development contracts. The GE design was carried forward into Integrated Core/Low Spool (ICLS) 

testing (Ref. 10). This full-system, integrated testing provides a full-engine dataset that was not available 

with the Pratt and Whitney design. The GE version was therefore selected for this study. Figure 1 

illustrates the GE EEE design. 

At the start of the NPSS program, one of the goals was to create a geometry database that could be 

freely distributed to participating stake-holders. A team of NASA engineers and contractors worked with 

the original CAD files (obtained from GE) and created a series of IGES surfaces representing the flow 

path and blade shapes for the EEE turbomachinery. These geometry files were used in a series of engine 

component and full system calculations by Hall (Ref. 4). 

The overall EEE turbomachinery geometry, is illustrated in Figure 2. This figure displays the various 

engine components and indicates that each blade shape is represented by a series of cross-sectional views. 

The fan is captured by approximately 20 cross-sections, while the LPT displays only 5 cross-sections. 

These cross-sections are stacked radially and the turning between each cross-section can be fairly large, as 

is seen in Figure 3. IGES surfaces were constructed to closely match the cross-sections, but the significant 

variation between cross-sections made a perfect fit challenging. The variation between the cross-sections 

and the IGES surface may be a source of potential error in the blade shape representations. 

NASA/TM—2015-218408 2 
Figure 1.—Energy Efficient Engine developed by General Electric.

Figure 2.—Engine component geometry illustrated as a series of blade cross-sections.

Figure 3.—Three-dimensional, perspective view of a stator blade showing 

the cross-sectional stacking and the radial shape variation.

HPC HPT LPT LPC

Bypass Duct

NASA/TM—2015-218408 3 
Geometry files published with this report are organized as listed in Appendix A. Each component 

series of blade shapes is organized into a separate directory such as HPC, HPT, LPT or LPC. End walls or 

the limits of the flow domain can be extracted from the overall geometry file: “EEE_engine_all.igs” or 

“hub and casing” curves are provided in each component subdirectory. Blade counts are captured in 

Appendix B. In ensemble, these files provide much of the basic information needed for turbomachinery 

simulation. Information of the film-cooling or bleed can be extracted from Holloway (Ref. 6), Timko 

(Ref. 7) and Bridgeman (Ref. 11). 

At the top directory, the “EEE_engine_all.igs” file is stored along with symmetric and unsymmetrical 

elements. The original tar file used to store all this data is also stored in the root directory. This original 

tar file stored the file information in a version-control application that mangled these file labels. However, 

for this publication, the file names have been repaired based on information contained within each IGES 

file. 

The quarter stage booster and fan are included in the LPC directory. A series of curves provide 

information on the splitter and part-span shroud. 

Two directories for the HPC can be seen. The first labeled “HPC” is derived from the same source as 

all the other geometry files (GE). The directory labeled “HPC-CR16558 Derived” was developed by 

Turner and Siddappaji at the University of Cincinnati, using only information in Holloway (Ref. 6). This 

reference noted that a variety of blade shapes were used in the HPC (NACA series, Circular Arcs and 

others), but it was unclear which blade shape was used for any particular rotor or stator. For this reason, 

the blade thickness was determined from a general area rule. These blade shapes display significant 

differences between the GE original CAD shapes and the IGES representations from Holloway (Ref. 6). 

Appendix C illustrates the difference in leading and trailing edge profiles between the two different series 

of files. Large discrepancies are seen in the rotors 8, 9 and 10 and several stators. 

The EEE HPC was developed in the 1980s with two different “builds” or test articles. The second 

build was developed to improve on the original geometry documented in Holloway (Ref. 6). Holloway 

represents the first build while the geometry GE CAD shapes represent the second build. The impact of 

this variation will be discussed in Section 3.0, Engine System Performance. 

HPT and LPT directories are organized as the other components. 

2.1 Open Data Sources for the EEE Geometry 

As has been noted in the previous section, the blade shapes for each component of the EEE can be 

largely reconstructed from the NASA Contractor Reports (Refs. 6, 7, 10, and 11). HPC shapes were 

extracted from Holloway (Ref. 6)for the first build of the EEE HPC. HPT blade coordinates were listed in 

Appendix A of Timko (Ref. 7). Scale drawings of the blade cross-section at three radial locations (hub, 

pitch and tip) are also available in this reference. Figure 1 in Timko provided detailed dimensions for the 

flow path. The LPT blade shapes were documented in Bridgeman (Ref. 11) through scale drawings of 

Figures 10 to 19. The flow path was documented in Figure 9. While these sources are not as reliable as 

the actual CAD representations, they do provide adequate information to analyze highly similar systems. 

3.0 Engine System Performance 

This section explores some numerical simulations of the EEE turbomachinery geometry documented 

in Section 2.0. New simulation results are placed in context by comparing with previous efforts (Ref. 4). 

Previous efforts were based on computer simulation technology of the 1990s with relatively course mesh 

resolution, constant gamma and a mixing plane treatment of the rotor / stator interaction. Modern CFD 

tools provide more options for numerical accuracy, much finer mesh resolution, and an improved 

treatment of the rotor / stator interaction using a Non-Linear Harmonic (NLH) method (Ref. 12). It was of 

interest to see if these improvements resulted in better system performance. 

Numerical simulations of the rotating turbomachinery involved solving the Reynolds’s Averaged 

Navier-Stokes equations using the Numeca FINE/Turbo code suite (Ref. 12). This software has a fully 

NASA/TM—2015-218408 4 
integrated suite of tools that enables rapid mesh generation, domain decomposition, flow solving, and 

visualization. The mesh generation can be controlled through a series of utilities that tailor the grid 

topology and optimization to the application-specific geometry. Typically, the defaults for the mesh 

generation provide good mesh characteristics; however, adjustments were made using available manual 

controls. FINE/Turbo employs a variety of possible rotor/stator interface techniques, turbulence models, 

and differencing techniques. In this study, the steady-state options were used, including the Mixing Plane 

approximation, the Spalart-Almos turbulence model, and Flux Difference Splitting (second order upwind) 

with Min Mod limiters. Some calculations of the HPC were made using the NLH method. 

Hall’s simulation results (Ref. 4) (representative of early modeling efforts) can be seen in Table 1 for 

the simulated EEE system values and a cycle model. This cycle model was matched to the experimental 

engine tests(Ref. 10) by Owen and documented in Hall (Ref. 4). As can be seen in Table 1, the turbines 

exhibit poor corrected mass flow, pressure ratios, and efficiencies. However, the HPC is very well 

matched and the other compression components are all within a few percent of the cycle model results. 

Given the state of mesh resolution, constant gamma and model complexity in this study, the overall 

predictive accuracy is surprisingly good. However, if all these component simulations were fully coupled 

to the 0–D NPSS model, it seems likely that the predicted engine performance would not closely resemble 

the EEE. 

With the recent advances in simulation capabilities, many of the EEE engine turbomachinery 

components have been analyzed again to better predict performance. Table 2 illustrates the component 

performance calculated using Fine/Turbo in these simulations for much of the EEE engine turbomachinery. 

The Fan/LPC component is not displayed as calculations of this component were not performed. Overall, 

the efficiency of each component is fairly good, with the exception of the LPT. The LPT displays a 

–3 percent performance variance. This level of inaccuracy may be due to a large number of factors, 

including poor development of the geometry, and model/mesh limitations. The good agreement seen in the 

HPT may be fortuitous as film cooling was not included and would likely alter the indicated performance. 

All calculations were performed with smooth hub and casing surfaces, whereas, a more realistic 

representation would include hub leakages, turbine blade film cooling, secondary flow paths, and a 

capturing of the “button” geometry for the variable angle stators in the HPC. These calculations employed 

a mixing plane assumption for the rotor / stator interaction plane and better models may have improved 

the calculated results. The following sections discuss details of each component simulation. 

TABLE 1.—ENGINE SYSTEM PERFORMANCE RESULTS OF 

HALL’S SIMULATIONS COMPARED WITH A 0–D CYCLE 

MODEL BASELINED TO EXPERIMENTAL DATA 

Wc, 

% 

PR, 

% 

Efficiency, % 

FAN 3.23 –1.80 1.16 

LPC 0.59 –1.67 1.99 

HPC –0.29 –0.22 –0.30 

HPT –6.17 –6.11 –2.88 

LPT 10.11 11.49 –1.32 

TABLE 2.—COMPONENT SIMULATION VARIANCES 

FROM THE BASELINE 0–D NPSS MODEL 

Component W, 

% 

PR, 

% 

Efficiency, 

% 

HPC 3.45 –0.68 –0.89 

HPT 6.64 1.55 0.32 

LPT 8.32 2.32 –3.00 

NASA/TM—2015-218408 5 
3.1 High Pressure Compressor (HPC) 

The HPC was the most difficult component to simulate as might be expected from the high pressureratio compression performed by the system. The HPC, nominally, could perform at a pressure ratio of 25 

although the calculations performed here were limited to about 23. As was reported earlier, the blade 

shapes directly traceable to GE CAD drawings had significant differences from the shapes developed 

from Holloway (Ref. 6). All calculations performed using the GE CAD geometry did not converge. The 

calculations reported here used the shapes derived from Holloway (Ref. 6). 

Tip clearances were illustrated in Holloway (Ref. 6). Figure 4 is recreated from Figure 62 of this 

reference. The estimated line was presumably calculated using measured temperatures and pressure from 

rig testing. Two probe tip-clearance measurement locations were documented and match well within the 

estimated range of value. 

The baseline HPC calculation seen in Table 2 was performed using the mixing plane assumption with 

flux-difference splitting of the momentum equations. The use of flux-difference splitting was important as 

central differencing resulted in reduced efficiencies and ultimately, the central difference calculation 

would not converge. The flux-differencing splitting, combined with mixing-plane, resulted in the work 

loading seen in Figure 5. This figure illustrates the percent of the total temperature rise occurring across 

each stage. The first two stages are highly loaded with the remaining stages experiencing between 9 and 

10 percent temperature rise. 

All HPC calculations were performed using 15,462,827 mesh nodes, with rotor blades typically 

employing about 1 million nodes and the stators at a lower level of resolutions of about 600K nodes. 

Figure 4.—HPC tip clearances as estimated and measured from 

Holloway.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

1 2 3 4 5 6 7 8 9 10

Tip Clearance (mm)

HPC Stage

Estimated

Probe

NASA/TM—2015-218408 6 
Figure 5.—Stage distribution of temperature rise (work loading) for the 

design and numerically calculated results.

3.2 High Pressure Turbine (HPT) 

The HPT calculations were performed using Fine-Turbo with 3,828,020 nodes, central differencing, 

and with the mixing plane model for rotor / stator interactions. This simulation employed smooth 

sidewalls and no cooling bleed. The cooling bleed and sidewall treatment are obvious areas for 

improvement in this calculation. Despite these limitations the system performance is well predicted. 

Nominal tip clearance was documented by Timko (Ref. 7) as 41 mm. 

From Hall (Ref. 4): “GE engineers familiar with the actual test rig and EEE engine geometry 

recommended a one degree (open) reset of the LP turbine first stage vane.” This reset is maintained in 

these calculations. 

3.3 Low Pressure Turbine (LPT) 

The LPT calculation was performed with 12,410,092 nodes, central differencing and the mixing plane 

model. The calculated efficiency of 89.7 percent is lower that the cycle results, but this value is 

reasonable considering the limitations on the calculation. The boundary layer transition may be an area 

requiring further study / variation and, as with the HPT, bleed and sidewalls need to be improved. In 

general, the computed results are promising, but further study is suggested. 

3.4 Summary 

Overall, the component performance results suggest that a full coupling of the high-fidelity 

simulations with a 0D cycle simulation may be possible. Most component results are a fair match with 

system data. Much more simulation testing would be required to quantify the level of uncertainty, but the 

geometry provided by this report is certainly a good starting point for further refinement of the geometry 

and analysis features. 

9.0%

9.5%

10.0%

10.5%

11.0%

11.5%

12.0%

1 2 3 4 5 6 7 8 9 10

Stage Temperature Rise (Percent Total)

STAGE

Design Work Profile

Calculated

NASA/TM—2015-218408 7 
Figure 6.—Typical EEE engine system model output at cruise conditions.

4.0 EEE Engine Cycle 

A 0–D representation of the ICLS engine was developed to support the simulations of Hall (Ref. 4). A 

typical output for cruise is illustrated in Figure 6. The complete NPSS model is included in accompanying 

compact disk. This model may be helpful to future efforts building a fully-coupled engine simulation. 

5.0 Summary 

The intent of this paper was to document the EEE geometry that may provide a baseline for testing 

full-engine simulation techniques. It appears highly likely that some of the geometry provided here may 

not be an exact representation of the hot, operating geometry for all components. The blade shapes 

received from GE did not contain blade surfaces, and the NURBS representations created for the blades 

may have flaws. Simulation results indicate reasonable component performance, but the accuracy is still 

subject to a variety of uncertainties. There is a great need to add detailed hub leakages models, variable 

stator “button” geometries, film cooling and end wall flows. The openly published contractor reports 

provide much of the information needed to further refine this geometry. The refined simulations might 

result in significantly different performance. 

NASA/TM—2015-218408 8 
Appendix A.—Supplemental Information 

Table 3 provides a listing of the files contained on the supplemental DVD available with this report. 

TABLE 3.—LISTING OF FILES BY FOLDER CONTAINED IN SUPPLEMENTAL DVD 

File Name Date Modified Size (bytes) 

Directory: \EEE 

assem.igs 12/12/2012 28,436,480 

directory.txt 12/19/2012 0 

eee.tar.gz 12/12/2012 14,202,965 

EEE_3d_sym.igs 12/12/2012 426,465 

EEE_3d_uns.igs 12/12/2012 9,596,232 

EEE_engine_all.igs 12/12/2012 8,736,255 

HPC (Folder) 12/19/2012 

HPC-CR16558 Derived (Folder) 12/19/2012 

HPT (Folder) 12/19/2012 

LPC (Folder) 12/19/2012 

LPT (Folder) 12/19/2012 

Directory: \EEE\HPC 

case.igs 11/30/2009 41,310 

compr_igv.igs 12/12/2012 194,319 

compr_rotor1.igs 12/12/2012 275,400 

compr_rotor10.igs 12/12/2012 187,353 

compr_rotor2.igs 12/12/2012 275,886 

compr_rotor3.igs 12/12/2012 249,804 

compr_rotor4.igs 12/12/2012 241,380 

compr_rotor5.igs 12/12/2012 241,461 

compr_rotor6.igs 12/12/2012 219,510 

compr_rotor7.igs 12/13/2012 222,426 

compr_rotor8.igs 12/12/2012 217,161 

compr_rotor9.igs 12/12/2012 195,048 

compr_stator1.igs 12/12/2012 196,506 

compr_stator10.igs 12/12/2012 171,477 

compr_stator2.igs 12/12/2012 184,032 

compr_stator3.igs 12/12/2012 146,448 

compr_stator4.igs 12/12/2012 139,806 

compr_stator5.igs 12/12/2012 118,422 

compr_stator6.igs 12/12/2012 107,730 

compr_stator7.igs 12/12/2012 119,637 

compr_stator8.igs 12/12/2012 166,698 

compr_stator9.igs 12/12/2012 148,716 

hub.igs 11/30/2009 41,310 

Directory: \EEE\HPC-CR16558 Derived 

blade1.igs 12/6/2011 376,083 

blade10_cf.igs 2/24/2012 311,928 

blade11_cf.igs 2/24/2012 308,730 

blade12.igs 12/6/2011 376,083 

blade13_cf.igs 2/24/2012 304,958 

blade14.igs 12/6/2011 376,083 

blade15.igs 12/6/2011 376,083 

blade16.igs 12/6/2011 376,083 

blade17.igs 12/6/2011 376,083 

blade18_cf.igs 3/23/2012 306,352 

blade19_cf.igs 4/4/2012 312,256 

NASA/TM—2015-218408 9 
TABLE 3.—LISTING OF FILES BY FOLDER CONTAINED IN SUPPLEMENTAL DVD 

File Name Date Modified Size (bytes) 

blade2.igs 12/6/2011 376,083 

blade20_cf.igs 4/4/2012 308,402 

blade21_cf.igs 4/4/2012 311,354 

blade3.igs 12/6/2011 343,845 

blade4.igs 12/6/2011 376,083 

blade5_cf.igs 2/24/2012 306,844 

blade6.igs 12/6/2011 376,083 

blade7.igs 12/6/2011 376,083 

blade8.igs 12/6/2011 376,083 

blade9.igs 12/6/2011 376,083 

case.igs 11/30/2009 41,310 

hub.igs 11/30/2009 41,310 

README.txt 12/19/2012 593 

Directory: \EEE\HPT 

case.igs 3/17/2010 21,402 

hpt_rotor1.igs 12/12/2012 118,422 

hpt_rotor2.igs 12/12/2012 113,967 

hpt_stator1.igs 12/12/2012 116,640 

hpt_stator2.igs 12/12/2012 107,568 

hub.igs 3/17/2010 21,402 

Directory: \EEE\LPC 

booster_rtr.igs 12/12/2012 270,135 

booster_stator.igs 12/12/2012 607,338 

Bypass_Outer.igs 3/23/2010 34,932 

bypass_stator.igs 12/12/2012 644,112 

core_guide_vane.igs 12/12/2012 402,327 

fan.igs 12/14/2012 944,541 

hub.igs 3/16/2010 22,468 

LPC_bypass_hub.igs 3/25/2010 14,094 

part_span_shroud_lower.igs 3/16/2010 6,068 

part_span_shroud_upper.igs 3/16/2010 6,068 

splitter_lower.igs 3/16/2010 21,238 

splitter_upper.igs 3/16/2010 21,238 

Directory: \EEE\LPT 

case.igs 3/17/2010 50,102 

hub.igs 3/17/2010 50,102 

lpt_rotor1.igs 12/12/2012 104,490 

lpt_rotor2.igs 12/12/2012 103,761 

lpt_rotor3.igs 12/12/2012 101,736 

lpt_rotor4.igs 12/12/2012 103,923 

lpt_rotor5.igs 12/12/2012 103,923 

lpt_stator1.igs 12/12/2012 98,253 

lpt_stator2.igs 12/12/2012 100,926 

lpt_stator3.igs 12/12/2012 99,954 

lpt_stator4.igs 12/14/2012 97,848 

lpt_stator5.igs 12/12/2012 94,932 

 

NASA/TM—2015-218408 10 
Appendix B 

Table 4 lists circumferential blade counts for the various turbomachinery components of the EEE. 

TABLE 4.—CIRCUMFRENTIAL BLADE COUNT FOR EACH BLADE ROW 

LPC 

Blade 

Count HPC 

Blade 

Count HPT 

Blade 

Count LPT 

Blade 

Count 

1 32 1 32 1 46 1 72 

2 60 2 28 2 76 2 120 

3 56 3 50 3 48 3 102 

4 64 4 38 4 70 4 122 

5 34 5 50 5 96 

 6 68 6 122 

 7 82 7 114 

 8 60 8 156 

 9 92 9 120 

 10 70 10 110 

 11 110 

 12 80 

 13 120 

 14 82 

 15 112 

 16 84 

 17 104 

 18 88 

 19 118 

 20 96 

 21 140 

 

NASA/TM—2015-218408 11 

Appendix C 

Figure 7 to Figure 28 document the difference between geometry-specified leading and trailing edges 

for all blades of the HPC. The blades labeled UC were developed by Turner and Kiran. The Allison 

shapes trace back to the original CAD drawings. 

 

Figure 7.—Comparison of the trailing and leading edge of the IGV. 

 

-10

-8

-6

-4

-2

0

2

4

6

8

10

012

Angle

% span

IGV Trailing Edge

 UC Trailing

Edge Angle

Allison Leading

Edge Angle

-10

-8

-6

-4

-2

0

2

4

6

8

10

012

Angle

% span

IGV Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

NASA/TM—2015-218408 13 
 

Figure 8.—Comparison of the trailing and leading edge of Rotor 1. 

 

Figure 9.—Comparison of the trailing and leading edge of Stator 1. 

 

30

35

40

45

50

55

60

65

70

012

Angle

% span

Rotor 1 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-20

-10

0

10

20

30

40

50

60

70

012

Angle

% span

Rotor 1 Trailing Edge

 UC Trailing

Edge Angle

Allison Leading

Edge Angle

-55

-50

-45

-40

-35

-30

-25

-20

012

Angle

% span

Stator 1 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-15

-10

-5

0

5

10

012

Angle

% span

Stator 1 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 14 
 

Figure 10.—Comparison of the trailing and leading edge of Rotor 2. 

 

Figure 11.—Comparison of the trailing and leading edge of Stator 2. 

 

30

35

40

45

50

55

60

65

70

012

Angle

% span

Rotor 2 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-10

0

10

20

30

40

50

60

012

Angle

% span

Rotor 2 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 2 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-25

-23

-21

-19

-17

-15

-13

-11

-9

-7

-5

012

Angle

% span

Stator 2 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 15 
 

Figure 12.—Comparison of the trailing and leading edge of Rotor 3. 

 

Figure 13.—Comparison of the trailing and leading edge of Stator 3. 

 

40

45

50

55

60

65

70

012

Angle

% span

Rotor 3 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

0

10

20

30

40

50

60

012

Angle

% span

Rotor 3 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-30

-25

-20

-15

-10

-5

0

012

Angle

% span

Stator 3 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-60

-55

-50

-45

-40

-35

012

Angle

% span

Stator 3 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

NASA/TM—2015-218408 16 
 

Figure 14.—Comparison of the trailing and leading edge of Rotor 4. 

 

Figure 15.—Comparison of the trailing and leading edge of Rotor 4. 

 

40

45

50

55

60

65

70

012

Angle

% span

Rotor 4 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

0

5

10

15

20

25

30

35

40

45

50

012

Angle

% span

Rotor 4 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-58

-56

-54

-52

-50

-48

-46

-44

-42

-40

012

Angle

% span

Stator 4 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-30

-25

-20

-15

-10

-5

012

Angle

% span

Stator 4 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 17 
 

Figure 16.—Comparison of the trailing and leading edge of Rotor 5. 

 

Figure 17.—Comparison of the trailing and leading edge of Stator 5. 

 

49

51

53

55

57

59

61

63

65

012

Angle

% span

Rotor 5 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

10

15

20

25

30

35

40

45

50

012

Angle

% span

Rotor 5 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-60

-58

-56

-54

-52

-50

-48

-46

-44

-42

-40

012

Angle

% span

Stator 5 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-30

-25

-20

-15

-10

-5

012

Angle

% span

Stator 5 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 18 
 

Figure 18.—Comparison of the trailing and leading edge of Rotor 6. 

 

Figure 19.—Comparison of the trailing and leading edge of Stator 6. 

 

52

54

56

58

60

62

64

66

012

Angle

% span

Rotor 6 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

15

20

25

30

35

40

45

012

Angle

% span

Rotor 6 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 6 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-18

-16

-14

-12

-10

-8

-6

-4

-2

0

012

Angle

% span

Stator 6 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 19 
 

Figure 20.—Comparison of the trailing and leading edge of Rotor 7. 

 

Figure 21.—Comparison of the trailing and leading edge of Stator 7. 

 

55

56

57

58

59

60

61

62

63

64

65

012

Angle

% span

Rotor 7 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

0

5

10

15

20

25

30

35

40

45

012

Angle

% span

Rotor 7 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-65

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 7 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-18

-16

-14

-12

-10

-8

-6

-4

-2

0

012

Angle

% span

Stator 7 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 20 
 

Figure 22.—Comparison of the trailing and leading edge of Rotor 8. 

 

Figure 23.—Comparison of the trailing and leading edge of Stator 8. 

 

56

57

58

59

60

61

62

63

64

65

012

Angle

% span

Rotor 8 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

15

20

25

30

35

40

45

012

Angle

% span

Rotor 8 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-65

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 8 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-18

-16

-14

-12

-10

-8

-6

-4

-2

0

012

Angle

% span

Stator 8 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 21 
 

Figure 24.—Comparison of the trailing and leading edge of Rotor 9. 

 

Figure 25.—Comparison of the trailing and leading edge of Stator 9. 

 

58

59

60

61

62

63

64

65

66

67

68

69

012

Angle

% span

Rotor 9 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

15

20

25

30

35

40

45

012

Angle

% span

Rotor 9 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-65

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 9 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

-16

-14

-12

-10

-8

-6

-4

-2

0

012

Angle

% span

Stator 9 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 22 
 

Figure 26.—Comparison of the trailing and leading edge of Rotor 10. 

 

Figure 27.—Comparison of the trailing and leading edge of Stator 10. 

 

60

61

62

63

64

65

66

67

68

69

012

Angle

% span

Rotor 10 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

20

25

30

35

40

45

012

Angle

% span

Rotor 10 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

-70

-65

-60

-55

-50

-45

-40

-35

-30

012

Angle

% span

Stator 10 Leading Edge

 UC Leading

Edge Angle

Allison Leading

Edge Angle

0

2

4

6

8

10

12

012

Angle

% span

Stator 10 Trailing Edge

 UC Trailing

Edge Angle

 Allison Trailing

Edge Angle

NASA/TM—2015-218408 23 
References 

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

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2010-475, Sept. 2010. 

3. Turner, M.G., Reed, J.A., Ryder, R. and Veres, J.P., “Multi-Fidelity Simulation of a Turbofan 

Engine with Results Zoomed into Mini-Maps for a Zero-D Cycle Simulation,” ASME Paper No. 

GT2004-53956, Vienna, Austria, June 2004 (also NASA/TM—2004-213076). 

4. Hall, E., Owen, P., “Energy Efficient Engine HP/LP Spool Analysis,” To Be Published. 

5. Cline, S.J., Kutney, J.T., Halter, P.H., Sulivan, T.J., “Energy Efficient Engine Fan and QuarterStage Component Performance Report,” NASA CR-168070, Jan. 1983. 

6. Holloway, P.R., Knight, G.L., Koch, C.C., Shaffer, S.J., “Energy Efficient Engine High Pressure 

Compressor Detail Design Report,” NASA CR-165558, May 1982. 

7. Timko, L.P., “ Energy Efficient Engine High Pressure Turbine Component Test Performance 

Report,” NASA CR-168289. 

8. Hall, E., Delaney, R., Lynn, S., and Veres, J. “Energy Efficient Engine Low Pressure Subsystem 

Aerodynamic Analysis,” NASA/TM—1998-208402. 

9. Hall, E., “Modular Multi-Fidelity Simulation Methodology for Multiple Spool Turbofan Engines,” 

NASA High Performance Computing and Communications Computational Aerosciences Workshop, 

NASA Ames Research Center, February 15-17, 2000. 

10. Stearns, EM, et al, “Energy Efficient Engine Integrated Core/Low Spool Design and Performance 

Report,” NASA CR-168211, 1985. 

11. Bridgeman, M.J., Cherry, D.G., Pedersen, J., “NASA /GE Energy Efficient Engine Low Pressure 

Turbine Scaled Test Vehicle Performance Test,” NASA CR-168290, July 1983. 

12. Vilmin, S., Lorrain, E., Hirsch, CH., Swoboda, M., “Unsteady Flow Modeling Across the Rotor / 

Stator Interface Using the Nonlinear Harmonic Method,” ASME GT2006-90210. 

NASA/TM—2015-218408 24