Geometry and Simulation Results for a Gas Turbine Representative of the Energy Efficient Engine (EEE)

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 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 3.1 High Pressure Compressor (HPC) The HPC was the most difficult component to simulate as might be expected from the high pressure- ratio 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.

0.7 0.6 0.5 0.4 Estimated 0.3 Probe 0.2 Tip Clearance (mm) 0.1 1 2 3 4 5 6 7 8 9 10 HPC Stage Figure 4.—HPC tip clearances as estimated and measured from Holloway.

NASA/TM—2015-218408 12.0% Design Work Profile Calculated 11.5% 11.0% 10.5% 10.0% 9.5% Stage Temperature Rise (Percent Total) 9.0% 1 2 3 4 5 6 7 8 9 10 STAGE 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.

NASA/TM—2015-218408 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

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 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

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 Count HPC Count HPT Count LPT Count Blade Blade Blade Blade 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

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.

IGV Trailing Edge IGV Leading Edge UC Trailing 2 UC Leading Edge Angle Edge Angle Angle Allison Leading Allison Leading -2 -2 Angle Edge Angle Edge Angle -4 -4 -6 -6 -8 -8 -10 -10 0 1 2 0 1 2 % span % span Figure 7.—Comparison of the trailing and leading edge of the IGV.

NASA/TM—2015-218408 Rotor 1 Trailing Edge Rotor 1 Leading Edge UC Leading UC Trailing Edge Angle Edge Angle Angle Angle Allison Leading Allison Leading Edge Angle Edge Angle -10 -20 0 1 2 0 1 2 % span % span Figure 8.—Comparison of the trailing and leading edge of Rotor 1.

Stator 1 Leading Edge Stator 1 Trailing Edge -20 -25 -30 UC Leading 0 UC Trailing -35 Edge Angle Edge Angle Angle Angle Allison Trailing -40 Allison Leading -5 Edge Angle Edge Angle -45 -10 -50 -15 -55 0 1 2 0 1 2 % span % span Figure 9.—Comparison of the trailing and leading edge of Stator 1.

NASA/TM—2015-218408 Rotor 2 Leading Edge Rotor 2 Trailing Edge UC Trailing UC Leading 30 Edge Angle Edge Angle 20 Angle Angle Allison Leading Allison Trailing Edge Angle Edge Angle -10 0 1 2 0 1 2 % span % span Figure 10.—Comparison of the trailing and leading edge of Rotor 2.

Stator 2 Leading Edge Stator 2 Trailing Edge -5 -30 -7 -35 -9 -11 -40 UC Leading -13 UC Trailing Edge Angle Edge Angle -45 -15 Allison Leading Angle Angle Allison Trailing -17 Edge Angle Edge Angle -50 -19 -21 -55 -23 -60 -25 0 1 2 0 1 2 % span % span Figure 11.—Comparison of the trailing and leading edge of Stator 2.

NASA/TM—2015-218408 Rotor 3 Leading Edge Rotor 3 Trailing Edge UC Trailing UC Leading 30 Edge Angle Edge Angle Angle Angle Allison Trailing Allison Leading Edge Angle Edge Angle 0 1 2 0 1 2 % span % span Figure 12.—Comparison of the trailing and leading edge of Rotor 3.

Stator 3 Leading Edge Stator 3 Trailing Edge -35 -5 -40 -10 UC Leading UC Trailing -45 Edge Angle Edge Angle -15 Allison Leading Angle Allison Trailing Angle -50 Edge Angle Edge Angle -20 -55 -25 -60 -30 0 1 2 0 1 2 % span % span Figure 13.—Comparison of the trailing and leading edge of Stator 3.

NASA/TM—2015-218408 Rotor 4 Leading Edge Rotor 4 Trailing Edge UC Leading UC Trailing Edge Angle 25 55 Edge Angle Angle Allison Leading 20 Angle Allison Trailing Edge Angle Edge Angle 0 1 2 0 1 2 % span % span Figure 14.—Comparison of the trailing and leading edge of Rotor 4.

Stator 4 Leading Edge Stator 4 Trailing Edge -40 -5 -42 -10 -44 -46 -15 UC Leading UC Trailing -48 Edge Angle Edge Angle -50 Angle Allison Leading Angle Allison Trailing -20 Edge Angle Edge Angle -52 -54 -25 -56 -58 -30 0 1 2 0 1 2 % span % span Figure 15.—Comparison of the trailing and leading edge of Rotor 4.

NASA/TM—2015-218408 Rotor 5 Leading Edge Rotor 5 Trailing Edge 65 50 63 45 59 35 UC Leading UC Trailing Edge Angle Edge Angle Allison Trailing Angle Angle Allison Leading Edge Angle Edge Angle 0 1 2 0 1 2 % span % span Figure 16.—Comparison of the trailing and leading edge of Rotor 5.

Stator 5 Leading Edge Stator 5 Trailing Edge -40 -5 -42 -44 -10 -46 -48 UC Leading -15 UC Trailing Edge Angle Edge Angle -50 Angle Allison Trailing Allison Leading Angle -52 -20 Edge Angle Edge Angle -54 -56 -25 -58 -60 -30 0 1 2 0 1 2 % span % span Figure 17.—Comparison of the trailing and leading edge of Stator 5.

NASA/TM—2015-218408 Rotor 6 Leading Edge Rotor 6 Trailing Edge 66 45 UC Leading UC Trailing Edge Angle Edge Angle 58 Allison Leading Allison Trailing Angle Angle Edge Angle Edge Angle 52 15 0 1 2 0 1 2 % span % span Figure 18.—Comparison of the trailing and leading edge of Rotor 6.

Stator 6 Leading Edge Stator 6 Trailing Edge -30 -2 -35 -4 -6 -40 UC Leading UC Trailing Edge Angle -8 Edge Angle -45 Allison Leading -10 Allison Trailing Angle Angle Edge Angle Edge Angle -12 -50 -14 -55 -16 -18 -60 0 1 2 0 1 2 % span % span Figure 19.—Comparison of the trailing and leading edge of Stator 6.

NASA/TM—2015-218408 Rotor 7 Leading Edge Rotor 7 Trailing Edge UC Leading UC Trailing Edge Angle Edge Angle Angle Angle Allison Trailing Allison Leading Edge Angle Edge Angle 15 57 10 0 1 2 0 1 2 % span % span Figure 20.—Comparison of the trailing and leading edge of Rotor 7.

Stator 7 Leading Edge Stator 7 Trailing Edge -30 -2 -35 -4 -40 -6 UC Leading -45 UC Trailing -8 Edge Angle Edge Angle -10 -50 Angle Angle Allison Leading Allison Trailing Edge Angle Edge Angle -12 -55 -14 -60 -16 -65 -18 0 1 2 0 1 2 % span % span Figure 21.—Comparison of the trailing and leading edge of Stator 7.

NASA/TM—2015-218408 Rotor 8 Leading Edge Rotor 8 Trailing Edge 65 45 UC Leading UC Trailing Edge Angle Edge Angle Angle Angle Allison Leading Allison Trailing Edge Angle Edge Angle 59 25 0 1 2 0 1 2 % span % span Figure 22.—Comparison of the trailing and leading edge of Rotor 8.

Stator 8 Leading Edge Stator 8 Trailing Edge -30 0 -2 -35 -4 -40 -6 -45 -8 UC Trailing UC Leading Edge Angle Edge Angle -50 -10 Angle Angle Allison Trailing Allison Leading Edge Angle -12 Edge Angle -55 -14 -60 -16 -65 -18 0 1 2 0 1 2 % span % span Figure 23.—Comparison of the trailing and leading edge of Stator 8.

NASA/TM—2015-218408 Rotor 9 Leading Edge Rotor 9 Trailing Edge 69 45 UC Leading UC Trailing Edge Angle Edge Angle Angle Allison Leading Allison Trailing Angle Edge Angle Edge Angle 58 15 0 1 2 0 1 2 % span % span Figure 24.—Comparison of the trailing and leading edge of Rotor 9.

Stator 9 Leading Edge Stator 9 Trailing Edge -30 0 -2 -35 -4 -40 -6 UC Leading UC Trailing -45 Edge Angle Edge Angle -8 Allison Leading Allison Trailing -50 Angle Angle Edge Angle -10 Edge Angle -55 -12 -60 -14 -16 -65 0 1 2 0 1 2 % span % span Figure 25.—Comparison of the trailing and leading edge of Stator 9.

NASA/TM—2015-218408 Rotor 10 Leading Edge Rotor 10 Trailing Edge UC Leading UC Trailing Edge Angle Edge Angle Angle Allison Leading Angle Allison Trailing Edge Angle Edge Angle 60 20 0 1 2 0 1 2 % span % span Figure 26.—Comparison of the trailing and leading edge of Rotor 10.

Stator 10 Leading Edge Stator 10 Trailing Edge -30 -35 -40 -45 UC Leading UC Trailing Edge Angle -50 Edge Angle Angle Allison Leading Allison Trailing Angle -55 Edge Angle Edge Angle -60 -65 -70 0 1 2 0 1 2 % span % span Figure 27.—Comparison of the trailing and leading edge of Stator 10.

NASA/TM—2015-218408

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