Edward P. Braunscheidel and Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Gary J. Skoch

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Gorazd Medic and Om P. Sharma

United Technologies Research Center, East Hartford, Connecticut

Aerodynamic Performance of a Compact,

High Work-Factor Centrifugal Compressor 

at the Stage and Subcomponent Level

NASA/TM—2015-218455

September 2015

AIAA–2014–3632
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NASA Langley Research Center

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Edward P. Braunscheidel and Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Gary J. Skoch

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Gorazd Medic and Om P. Sharma

United Technologies Research Center, East Hartford, Connecticut

Aerodynamic Performance of a Compact,

High Work-Factor Centrifugal Compressor 

at the Stage and Subcomponent Level

NASA/TM—2015-218455

September 2015

AIAA–2014–3632

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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cosponsored by the AIAA, ASME, SAE, and ASEE

Cleveland, Ohio, July 28–30, 2014
Acknowledgments

The effort was conducted under NASA Research Announcement cost-share contract funded by the NASA Fundamental Aeronautics 

Program, Rotary Wing Project, and the United Technologies Research Center (UTRC). Contributions from UTRC, including 

Drs. Joo Jongwook, Larry Hardin, Duane McCormick, William Cousins, Brian Holley, Paul Van Slooten, Elizabeth Lurie, 

Aamir Shabbir, and at NASA Glenn Research Center (GRC), including Dr. Michael Hathaway and Mark Stevens, are gratefully 

acknowledged. The efforts of the GRC test cell staff including Jonathon Mitchell, Jozsef Puskas, David Haydu, Adam Redding, 

and Richard Senyitko are greatly appreciated.

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at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2015-218455 1

Aerodynamic Performance of a Compact, High Work-Factor 

Centrifugal Compressor at the Stage and Subcomponent Level

Edward P. Braunscheidel and Gerard E. Welch

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Gary J. Skoch

U.S. Army Research Laboratory

Glenn Research Center

Cleveland, Ohio 44135

Gorazd Medic and Om P. Sharma

United Technologies Research Center

East Hartford, Connecticut 06118

Abstract

The measured aerodynamic performance of a compact, high work-factor, single-stage centrifugal 

compressor, comprising an impeller, diffuser, 90°-bend, and exit guide vane is reported. Performance 

levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data 

acquired at key machine rating planes during recent testing at NASA Glenn Research Center. 

Aerodynamic performance at the stage level is reported for operation between 70 to 105 percent of design 

corrected speed, with subcomponent (impeller, diffuser, and exit-guide-vane) flow field measurements 

presented and discussed at the 100 percent design-speed condition. Individual component losses from 

measurements are compared with pre-test CFD predictions on a limited basis. 

Nomenclature

D impeller trailing edge diameter (in.)

M Mach number

Nc corrected speed (rpm)

R radius (in.)

Re 



 





μ

ρ

in

inU2D2 , Reynolds number

SM ( )

( )

( )

( ) 















⋅ − ⋅

ref c

stall c

stall

ref

m

m

p p

p p





0,3 0,0

0,3 0,0

/

/

100% 1 , stability margin (NASA definition) 

T0 total temperature 

U rΩ = tip speed (ft/s)

b impeller trailing edge blade height (in.)

p0 total pressure

p static pressure

ṁ mass flow rate (lbm/s)

n exponent of Reynolds number correlation

u (ur, uθ, uz), velocity (ft/s)
NASA/TM—2015-218455 2

ε impeller tip-gap (in.)

ηtt total-to-total adiabatic efficiency

ηp,tt total-to-total polytropic efficiency

ηts total-to-static adiabatic efficiency

α atan(uθ/ur) or atan(uθ/uz), yaw angle in degrees

Ω angular frequency, (rad/s)

Subscripts / Stations

0, in inlet plenum

1 impeller inlet

2 impeller exit

2.2 probe survey location in vaneless space

2.4 diffuser main blade leading edge

2.5 diffuser main-to-main blade throat

2.6 splitter blade leading edge

2.7 diffuser exit

2.8 EGV inlet / leading edge

3, ex EGV exit / exit rating plane

PS pressure side

SS suction side

c corrected condition

max denotes maximum diameter of shroud endwall in 90° bend

ref reference condition (Nc = 100 percent, ṁc,ex = 3.0 lbm/s, ε = 0.012 in.)

stall condition at stall / stability boundary

1.0 Introduction

High pressure compressors of advanced turboshaft engines for rotorcraft applications comprise 

centrifugal compressor aft stages that achieve efficient pressure rise at low corrected flows and compact 

size. A recent collaborative effort between NASA and the United Technology Research Center (UTRC) 

related to a new centrifugal compressor—the “High Efficiency Centrifugal Compressor (HECC)”—is 

reported on herein.

The work, conducted under a NASA Research Announcement (NRA) cost-share contract, had intent 

to identify key technical barriers to advancement of aerodynamic performance of low corrected-flow 

(e.g., ṁc,ex  0.75 lbm/s engine scale) centrifugal compressor aft stages, and to acquire research-quality 

flow field data needed to close associated knowledge/data gaps. An aggressive aerodynamic goal set was 

established by NASA (Ref. 1) for a new state-of-the-art (SOA) centrifugal compressor. The test article 

was designed, fabricated and instrumented by UTRC (Refs. 2 and 3) under the NRA, and subsequently 

installed and tested at NASA Glenn Research Center (GRC). If achieved, design-intent aerodynamic 

performance levels would advance compressor efficiency (> 0.88 ηp,tt) levels at exit-corrected flow, work 

factor, stability margin, and diffuser compactness levels needed for advanced turboshaft engines (Ref. 1). 

UTRC recently documented the NRA effort in a NASA Contractor Report (Ref. 3), which included a 

detailed assessment of design-intent performance relative to the goal set. The current paper, focused on 

the experimental rating measurements, is intended to document the HECC performance at the stage

(overall map) and subcomponent (impeller, diffuser, 90° bend, and exit guide vane row) levels, with a 

brief investigation of performance derivatives associated with tip-clearance and Reynolds number 
NASA/TM—2015-218455 3

variations. Reported performance levels are based on steady-state total-pressure and total-temperature 

rake and total-pressure angularity-probe data acquired at key machine rating planes during testing at the 

GRC. Aerodynamic performance at the stage level is reported for operation between 70 to 105 percent of 

design corrected speed, with subcomponent (impeller, diffuser, and exit-guide-vane) performance 

discussions focusing mainly at 100 percent of design condition.

The requirements and aerodynamic design parameters of the HECC are first described, followed by a 

description of the experiment, including the test facility, data acquisition system, instrumentation and 

measurements. Finally a brief discussion related to comparison of design intent and achieved compressor 

performance is provided.

2.0 High Efficiency Centrifugal Compressor Description

2.1 Requirements

The challenging goal-set for the NRA solicitation of 2008 required high compressor efficiency 

(> 0.88 ηp,tt) at levels of exit-corrected flow, work factor, stability margin, and diffuser compactness 

(Dmax/D2) required for advance turboshaft engines for rotorcraft (Ref. 1). In addition to performance 

requirements, the NRA test article would be sized to enable research data acquisition and compatibility

with the existing running gear and facility interfaces of the Small Engine Component Test Facility (CE18) (Ref. 4) of GRC, where testing was to be conducted. The resultant UTRC test article was designed

(Refs. 2 and 3) to leverage existing NASA components – the impeller tip diameter, rotational speed, and 

inlet hub and shroud radii were fixed at values of the NASA CC3 compressor (Ref. 5). The goal-set 

corresponding to the 2x engine scale HECC test article is summarized in Table 1.

2.2 HECC Stage Description

The HECC stage shown schematically in Figure 1 comprises a splittered impeller, splittered diffuser, 

90° bend, and EGV blade row. The impeller has 15 main-blade/splitter-blade pairs with a spanwise 

varying backsweep between 32° to 42° from radial, and with –2° axial lean at the leading edge and –29°

axial lean at the trailing edge. The impeller blades have elliptical leading edges. Trailing edges are also 

elliptical which differs from the typically blunter trailing edges of trimmed impellers. The vaned diffuser 

comprises 20 main/splitter vane pairs with the splitters slightly offset to maximize pressure recovery. The 

EGV blade row comprises 60 cascade style airfoils. The design speed is Nc = 21,789 rpm, corrected to 

inlet conditions, resulting in an impeller exit speed of U2 = 1615 ft/s. Key physical features of the 

compressor are provided in Table 2. A comprehensive summary of the HECC aerodynamic and 

mechanical design effort is provided in References 2 and 3. 

TABLE 1.—GOAL SET FOR 2X-SCALE HECC COMPRESSOR

Metric Intent at rig scale

Exit corrected flow rate, ṁc,ex 2.1 < ṁc,ex < 3.1 lbm/s

Work factor 0.60 < ∆H0/U2

2 < 0.75

Polytropic efficiency, ηp,tt, % ≥ 0.88

Diameter ratio Dmax/D2 ≤ 1.45

Stall Margin, SM, % 13

Exit Mach number, Mex 0.15

Exit flow angle, αex 15°
NASA/TM—2015-218455 4

Figure 1.—High Efficiency Centrifugal Compressor stage (Ref. 2).

TABLE 2.—HECC DESIGN SUMMARY

Value Hub Shroud

Impeller blade inlet diameter, in. 3.19 8.479

Impeller splitter inlet diameter, in. 5.315 9.474

Impeller exit diameter, in. 16.988 16.998

Impeller exit blade height, in. 0.609

Impeller leading edge lean angle, deg. –2

Impeller trailing edge lean angle, deg. –29

Number of blade/splitter pairs 15

Tip clearance, in. 0.012

Vaned diffuser inlet diameter, in. 18.309 18.309

Vaned diffuser exit diameter, in. 22.398 22.398

Vaned diffuser channel height, in. 0.559

Number of vane/splitter pairs 20

Deswirl (EGV) blade inlet diameter, in. 23.663 24.594

Deswirl (EGV) blade exit diameter, in. 23.737 24.61

Deswirl (EGV) vane axial chord, in. 2.438

3.0 Facility and Compressor Test Rig Description

HECC testing was conducted in the NASA Small Engine Components Test Facility (cell CE-18, 

Figure 2) (Ref. 4) located at GRC. The facility was operated with atmospheric intake and exhaust, and an 

orifice plate was used to determine the physical mass flow rate. Inlet valves located downstream of the 

orifice are used to set the plenum pressure. To minimize inlet flow field disturbances, the inlet piping has 

long radius elbows with turning vanes and the plenum tank contains a foreign object protection screen, 

downstream of which is the compressor inlet reference (Station 0).

The facility throttle valve consists of a circumferentially rotating sleeve valve driven by a hydroelectric actuator with a control system providing extremely fine control of compressor flow rates. A

6,000 hp variable frequency motor drives a dual gearbox configurable for output speeds up to 60,000 rpm.

Drive motor speed can held to ± 1 rpm, resulting in a compressor speed uncertainty of ± 17 rpm, due to 

the drivetrain overall gear ratio. 

Impeller

Diffuser

EGV

90° Bend
NASA/TM—2015-218455 5

Figure 2.—Small Engine Components Test Facility (GRC CE-18).

3.1 Tip-Clearance Variation Capability

The main shaft of the compressor rig, and hence the impeller, could be translated axially during 

operation. The clearance gaps between the impeller tip and the shroud were measured using capacitance 

probes. Tip clearances are referred to as “inlet,” “knee,” and “exit”: the inlet measurement is at the 

impeller (main blade) leading edge, the knee is at the splitter blade leading edge, and the exit is at the 

trailing edge of the impeller blades (Figure 3(a)). The probe axis is designed to be normal to the local 

blade tip geometry at design operating conditions. Circumferential tip-clearance variations based upon the 

rub probe measurements are detailed in Figure 3(b). These measurements are estimated to be within 

± 0.001 in. of the actual value. The clearance asymmetries were consistent at all speeds, and the data 

reported herein were not adjusted to account for these asymmetric clearances. The tip clearance value

used to establish test conditions, and for tabulating results, was the maximum value at the impeller 

trailing edge for that operating condition.

3.2 Hub Flow Path Alignment

The impeller and diffuser hub flow paths were designed to be line-to-line at 100 percent Nc and designintent tip clearance of ɛref = 0.012 in. The hub flow path at the impeller exit exhibits a forward facing step for 

ɛ > ɛref , and an aft facing step for ɛ < ɛref . The magnitude of the step is equal to (ɛ – ɛref). Due to the asymmetry 

of impeller to the shroud, as previously discussed (Figure 3(b)), an aft facing step of approximately 0.003 in. 

exists for 50 percent of the impeller exit circumference at the design point operating condition.

3.3 Backplate Leakage Flow

Backplate leakage flow was controlled through the use of buffer air injected into the center of a multirow labyrinth seal located at the base of the impeller (Figure 3(a)). Service air was injected into the center 

groove until the static pressure therein matches the static pressure measured on the impeller backplate. 

The back-plate pressure measurements were used to automatically control the buffer air system. The 

system was operated continuously during testing to minimize leakage flow through this interface.

Mass Flow Orifice

Atmospheric 

Inlet

Collector and 

Throttle valve

Plenum Tank

(Station 0)

Compressor

Inlet control

valves

Atmospheric 

Exhaust

Gearbox
NASA/TM—2015-218455 6

(a) (b)

Figure 3.—(a) Tip clearance probe locations and (b) HECC clearance variations (in inches)

4.0 Data Acquisition, Instrumentation, and Test Procedures

In this section, the steady-state measurements used for rating the compressor will be summarized, and 

detailed subcomponent measurement stations will be identified. Measurement stations with unique 

instrumentation will be described in greater detail. Brief summaries of the data sets collected and how the 

stage was operated will be provided, along with estimates of measurement uncertainty. 

4.1 Mass Flow Rate and Compressor Operating Speed

Compressor mass flow rate was measured by using an orifice (Figure 2), with calculations and 

installation based upon the ASME standards (Ref. 6). Uncertainty for the orifice based mass flow 

measurement system is estimated to be ± 0.1 lbm/s, and day to day repeatability was observed to be 

± 0.05 lbm/s. No corrections for humidity were applied to any of the reported data. The relative humidity 

ranged from 40 to 75 percent during the testing timeframe.

Compressor operating speeds are based on once-per-rev measurements acquired at the compressor 

shaft near the aft face of the impeller. Compressor speeds are corrected to standard day conditions based 

upon temperature measurements taken in the plenum tank (Station 0).

4.2 Instrumentation and Measurement Stations

Compressor measurement stations are identified in Figure 4. The measurements acquired at each 

station are summarized in Table 2. Flow path (rakes and wall static) pressures are collected using a 

scanning pressure measurement system with an uncertainty of ± 0.07 percent of the transducers full scale. 

This results in uncertainties for the inlet pressure measurements of ± 0.011 psi, and for exit rakes 

± 0.06 psi. Temperature measurements are referenced to an isothermal reference block with an estimated 

uncertainty of ± 2.3 R for inlet conditions, and ± 2.5 R for stage exit measurements. Uncertainties were 

based on 2-sigma, 95 percent confidence level.

The data system scans all measurement channels at a sampling rate of 1 Hz. Data reported in this 

paper are averages of 20 consecutive scans. Prior to collecting data, the rig is verified to have reached 

thermal equilibrium based on tip-clearance measurements that remain constant over a 15 min time period. 

Pressure settling time is much quicker than the thermal response of the rig. When small changes are made 

to the operating condition, a minimum dwell time of 2 min was observed prior to data collection.

0.0094

0.0114

0.0111

0.0095

0.0098

0.0118

0.0127 0.012 0.0096 0.0088 0.0090

0.0092

Impeller Exit

Knee - Splitter 

Leading Edge

Inlet - Impeller 

Leading Edge

Inlet

Knee

Exit

Shroud Impeller 

Backplate

Hub Labyrinth 

seal
NASA/TM—2015-218455 7

Figure 4.—Compressor Measurement Stations

Wall static pressure taps located throughout the machine were checked and verified as they were 

connected to the pressure measurement system. Details of pressure measurement locations can be found 

in Reference 3.

4.2.1 Station 0—Inlet Plenum 

The key machine-level (overall) rating planes are at Stations 0 and 3 (Figure 4). All performance 

measurements are referenced to the inlet plenum, Station 0, where redundant thermocouples are averaged 

to define the T0,0, and four pressure ports equally spaced around the plenum circumference are sampled 

and averaged to obtain the p0,0 inlet reference conditions.

4.2.2 Station 2.2—Impeller Exit Survey and Related Test Procedure

Station 2.2 vaneless space surveys, conducted at a radius ratio (r2.2/r2) of 1.047, were performed using 

a three-port cobra probe (Figure 5) which was traversed from shroud to hub and manually nulled to within 

± 2° of the actual flow angle based on minimizing the side port pressure differentials. The probe was 

calibrated in a steady laminar free jet over a range of Mach numbers and flow angles of 0.3 to 0.84 and

± 20°, respectively. Calibration coefficients to correct for flow angle misalignment based on side port 

pressures, and coefficients to correct for pressure measurements at the misaligned flow angle, were 

generated and applied based on the approach in Reference 7. Flow angle measurements are typically 

within ± 1.0° in the laminar jet, and pressure is within ± 0.1 percent of the measured jet total pressure.

When testing with the probe at Station 2.2, a special operational procedure was required. As the probe 

was traversed from shroud to hub, the additional flow path blockage back-pressured the impeller, which 

was observed as a decrease in the inlet corrected flow rate. If no throttle change was made, the 

compressor stage would eventually stall as the blockage increased with increased probe immersion. Since 

the intent of this measurement was to measure the impeller exit flow at a constant condition, the following 

approach was implemented. With probe retracted, the operating condition was established, setting inlet 

corrected flow rate and stage pressure ratio. As the probe was traversed from shroud to hub the throttle 

was adjusted to maintain the initial inlet corrected mass flow rate. Using this procedure, the static 

pressures at the impeller exit were constant as the throttle was changed, reflecting that the impeller

Station 0

Plenum, Inlet 

Reference 

Station 1

Impeller Inlet

Station 2

Impeller Exit

Station 2.4

Diffuser Main Blade 

Leading Edge

Station 2.7 

Diffuser Exit

Station 2.8 

EGV Inlet

Station 3

EGV Exit/Rating 

Plane

Station 2.6 

Diffuser Splitter 

Leading Edge

Station 2.2

Vaneless Space 

Survey
NASA/TM—2015-218455 8

Figure 5.—Three-port Cobra probe

Figure 6.—Leading edge Kiel head instrumentation on main diffuser vane

operated at the initial set point. The survey measurements were obtained at corrected speed of 85 percent 

≤ Nc ≤ 100 percent at the design exit-corrected mass flow rate of 3.0 lbm/s. When testing was performed 

with probes at Station 2.2, the diffuser exit probes were not installed.

4.2.3 Station 2.4—Leading Edge Instrumented Diffuser Vanes

Station 2.4 measurements were obtained from main diffuser blade leading edge Kiel-head totalpressure ports, located on two identical vanes. Seven pressure ports, on equally space area centroids, span 

the passage height. Each vane was calibrated for Mach numbers from 0.3 to the calibration facility limit 

of 0.75, at flow angles from ± 20°. Note that the expected Mach numbers at the diffuser leading edge are 

> 0.75. Results show that within a range of ± 10° the effect of incidence on the measured vane total 

pressure relative to the jet pressure is approximately ± 0.25 percent. No corrections to the vane pressure 

measurements were applied for this report. Details of the modular vane are shown in Figure 6. The 

instrumented vanes were installed for all test conditions reported herein.

45o 0.325”

0.02”

0.065”

0.08”

0.15”
NASA/TM—2015-218455 9

4.2.4 Station 2.7—Diffuser Exit Survey and Related Test Procedure

Diffuser exit measurements at Station 2.7 were obtained at a radius ratio of 1.368. Six access locations 

span one main-to-main diffuser vane pitch at approximately equal spacing (Figure 7(a)). Two separate 

measurement techniques were used at Station 2.7: i) Using the probe shown in Figure 5, testing was 

performed one port at a time at design speed for three throttle settings identified as choke, design, and near 

stall. The probe was nulled and traversed from hub to shroud. Measurements were normalized to the 

plenum, and then collected together to obtain a complete mapping of the diffuser exit flow field. ii) All ports 

were simultaneously sampled using six identical miniature cobra probes which had sensing head sizes that 

were 50 percent smaller than the probe previously described. The probes were manually set at 15, 25, 50, 

75, and 85 percent spanwise locations, with angles being visually aligned with a reference angle on the 

mounting plug. All data were normalized to the plenum, and then collected to obtain a complete mapping of 

the diffuser exit flow field. Speed lines from choke to near stall were obtained at 85, 90, 95, and 100 percent 

of design speed with the six probe technique. All probes were calibrated in a similar manner as previously 

described at a Mach number of 0.4, and for yaw angles from ± 20°. The accuracies of the yaw angle 

measurements and immersions are not as consistent as with the actuated probe due to the manual nature of 

the setup. The yaw angles are estimated to be ± 3° of the reported values. The immersion of the probes is 

estimated to be within ± 0.02 in. of the reported value, which corresponds to an uncertainty in position of 

± 3 percent of span at Station 2.7. The test hardware is shown in Figure 7(b). When testing was performed 

with probes at Station 2.7, probes at Station 2.2 were not installed. 

4.2.5 Station 2.8—EGV Leading Edge Rakes

Total-pressure measurements were obtained from three adjacent EGVs with Kiel head leading edge 

total-pressure ports. The pressure ports were located at 25, 50 and 75 percent of the inlet span. Three 

EGVs span one main-to-main diffuser passage. Although no EGV calibrations were performed, the totalpressure port design was consistent with that used for all leading edge instrumentation. This 

instrumentation was installed for all test conditions. 

(a) (b)

Figure 7.—(a) Solid model depiction of diffuser exit measurements and (b) Actual hardware

Passage B Passage A
NASA/TM—2015-218455 10

4.2.6 Station 3.0—Machine Exit Rating Rakes

The compressor exit rating plane was located at Station 3, approximately 0.13 EGV chords 

downstream of the EGV trailing edge. The 12 exit rakes have three Kiel head total-pressure elements, and 

two Kiel head total-temperature elements each. The pressure elements were located on area centroids at 

10, 50, 92 percent of exit passage height, and the temperature elements were also located on area 

centroids at 30 and 70 percent of exit height. The rakes were calibrated for Mach number and acceptance 

angle sensitivities. The acceptance angle of the rakes was greater than ± 20°, no correction was applied 

due to their large acceptance angle. The 12 rakes span one main blade to main blade diffuser passage, and 

were yawed at 15° from axial to align with the design intent exit swirl angle. The rakes were distributed 

around the exit circumference as shown in Figure 8. The rakes were installed for all test conditions. 

Table 3 provides a summary of all Stations with their respective measurements.

Figure 8.—Exit rake (Station 3) locations

TABLE 3.—STATION MEASUREMENT SUMMARY (FIGURE 4)

a

Station Description Static pressures p0 T0 Flow angle,

Hub S α hroud

0 Inlet/Plenum X X X

1 Impeller Inlet X X

2 Impeller Exit X X

2.2 Vaneless space X X O O

2.4 Diffuser Main Vane X X X

2.5 Diffuser throat X X

2.6 Diffuser Splitter Vane X X

2.7 Diffuser Exit X X O O

2.8 EGV Inlet / Leading Edge X X X

3 EGV Exit / Rake Rating Plane X X X X a

“X” denotes permanently installed instrumentation

“O” denotes optional probe type measurements.
NASA/TM—2015-218455 11

5.0 Aerodynamic Performance Test Results

Aerodynamic performance results are provided in this section in terms of overall stage performance, 

then impeller, diffuser, and EGV subcomponent performance. Performance data were acquired from 

choke to stall, Nc = 70, 75, 85, 90, 95, 100 and 105 percent of design speed and as close as possible to the 

design exit tip clearance of 0.012 in. 

Facility thrust load limitations required testing at a suppressed inlet pressure. All data were acquired 

with the plenum total-pressure (Station 0) set to approximately p0,0 = 11 psia. 

The data reduction process for obtaining the one-dimensional (1-D) values used in the performance 

calculations is first described, followed by discussion of measured stage and subcomponent level 

performance metrics.

5.1 Performance Calculations/Data Reduction

The 1-D pressures and temperatures at the rating-plane and related machine-level and subcomponentlevel metrics reported herein are based on area averages at the rake and annulus levels. A curve fit of gas 

tables was used for the efficiency calculations. Adiabatic and polytropic efficiencies were determined by 

comparing the ideal enthalpy rise associated with the measured pressure ratio and enthalpy rise calculated 

using T0 measurements from the rig. The approach accounts for temperature-dependent fluid property 

changes. As noted, the performance metrics were not humidity-corrected.

5.2 Stage Performance

Overall performance at design clearance for the compressor stage as measured at the p0,0 = 11 psia 

suppressed inlet pressure is presented in Figure 9 and Figure 10. The stability boundary and a reference 

exit-corrected mass flow condition (“throttle line”) of ṁc,ex = 3.0 lbm/s are shown. The data point defining 

the stability boundary corresponds to the last stable operating condition at that speed. The stability margin 

for all speeds is provided in Table 4. 

At the design condition—that is, ṁc,ex = 3.0 lbm/s exit corrected flow, tip speed of Uc,0 = 1615 ft/s, and a 

tip clearance of 0.012 in.—the total-to-total pressure ratio was p0,3/p0,0 = 4.68, with a total-to-total adiabatic 

efficiency of 82.2 percent, polytropic efficiency of 85.5 percent, and SM = 7.5 percent. The achieved stage 

performance data at the design condition is compared to design-intent performance corrected for p0,0 = 11 

psia suppressed-inlet conditions in Table 5. As shown, the compressor adiabatic efficiency (ηtt) is low by 

2.75 points, the inlet corrected flow is low by 2.25 percent, and the exit Mach number and swirl, inferred 

from the exit rating plane measurements and machine flow, are high. The stability margin of the machine is 

significantly (4.5 points) lower than SM = 12 percent, the rig-scale design intent.

In this section, individual component performance levels for the impeller, diffuser, 90° bend, and 

EGV are presented, both to illustrate how the stage is performing as function of operating condition and 

to provide supporting evidence for a determination of where the subcomponent design may be 

underperforming. 
NASA/TM—2015-218455 12

Figure 9.—Ratio of total pressure (p0,3/p0,0) as a function of inlet corrected flow at 

design clearance.

Figure 10.—Total-to-total adiabatic (ηtt, closed symbols) and polytropic efficiency (ηp,tt,

open symbols) as a function of inlet corrected flow at design clearance.

1.5

2.0

2.5

3.0

3.5

4.0

4.5

5.0

5.5

4 5 6 7 8 9 10 11 12 13

Total Pressure Ratio, 

p0,3.0/p0,0

Inlet Corrected Mass Flow Rate, lbm/s

Stall Boundary

3.0 lbm/s op line

Design Point

105%

100%

95%

90%

85%

75%

70%

0.6

0.65

0.7

0.75

0.8

0.85

0.9

0.95

4.0 5.0 6.0 7.0 8.0 9.0 10.0 11.0 12.0 13.0

Efficiency, % 

Inlet Corrected Mass Flow Rate, lbm/s

Adiabatic

Design Intent

Polytropic
NASA/TM—2015-218455 13

TABLE 4.—SURGE MARGINS

AT 0.012 in. CLEARANCE

% Speed Surge margin

105 6.4%

100 7.5%

95 8.2%

90 10.6%

85 9.3%

75 18.4%

70 22.4%

TABLE 5.—MEASURED DESIGN POINT PERFORMANCE COMPARED TO DESIGN-INTENT PERFORMANCE 

[Nc = 100 percent, 0.012 in. tip clearance, suppressed inlet p0,0 = 11 psia.]

Metric Design Goal Rig scale

design intent

p0,0=14.7 psia

Rig scale

design intent

p0,0=11 psia

Measured

p0,0=11 psia

(±uncertainty)

Pressure ratio, p0,3/p0,0 4.85 4.80 4.68 ± 0.0074

Inlet flow rate, ṁc,in, lbm/s 11.2 11.1 10.85 ± 0.1

Exit flow rate, ṁc,ex, lbm/s 2.1 < ṁc,ex < 3.1 2.98 2.98 2.98

Adiabatic efficiency , ηtt, % 0.862 0.8495 0.822 ± 0.011

Polytropic efficiency, ηp,tt, % > 0.88 0.888 0.879 0.855

Adiabatic, total pressure to static pressure, ηts, % 0.852 0.8396 0.805

Exit Mach number, Mex 0.15 0.15 0.15 0.18

Exit flow angle, αex, deg 15° 14° 14° 34.3°

Stability Margin, SM, % 13 12 12 7.5

Work factor 0.60 < ∆H0/U2

2 < 0.75 0.7905 0.793 0.81

Diameter ratio Dmax/D2 < 1.45 1.45 1.45 1.45

5.3 Impeller—Exit Survey Data and Performance

5.3.1 Spanwise Discharge

The impeller performance is estimated using the Station 2.4 total pressure measurements and 

Station 3 total temperature measurements. Station 2.2 three-hole probe survey data are acquired as well. A 

comparison of the spanwise profile of total pressure from the survey probe, instrumented vanes and pretest CFD (design-intent) is provided in Figure 11(a). Note that reported pre-test CFD was conducted at 

standard day, rather than suppressed-inlet, conditions. The diffuser instrumented vane data agree well and 

are in qualitative agreement with the probe data, but the probe pressures are lower than expected. While 

the Kiel headed vane instrumentation at Station 2.4 was able to resolve a total pressure field which had

yaw angle fluctuations, the three-port cobra probe only resolved the actual total pressure when the yaw 

angle misalignment was close to zero. Even when the probe was nulled, the pressure indicated by the 

probe center port provided an integrated (average) value that is lower than the local p0, potentially due to 

the flow angle fluctuations and other unsteady effects. Design-point impeller exit swirl angle, obtained 

using the three-hole probe at Station 2.2, is provided in Figure 12 along with the corresponding pre-test 

CFD result. Although in qualitative agreement, the design-intent turning achieved by the impeller is 

higher than measured. Due to the uncertainties in the cobra probe results, the probe measurements at 

Station 2.2 were only used for qualitative assessments; that is, their values were not used in detailed 

component performance analyses.

To assess the impact of speed on impeller discharge, the probe measurements were reduced by the 

respective spanwise average to obtain the collapse shown in Figure 11(b). Evidently the spanwise profile 
NASA/TM—2015-218455 14

of total-pressure is relatively independent (on a reduced basis) of speed over the range (85 percent ≤ Nc ≤ 

100 percent) considered, with the exception of the 85 percent Nc, which appears to be an outlier in this 

respect (cf. stall boundary in Figure 9).

(a)

(b)

Figure 11.—(a) Comparison of design-point total-pressure at impeller exit. (Stations 2.2 and 2.4) 

and pre-test CFD; and, (b) Comparison of total-pressure coefficient measured at impeller exit 

as a function of speed.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

3 3.5 4 4.5 5 5.5 6

Fraction of span from hub

Local p0,2.2(z)/p0,0

NRA pre-test CFD

Modular Vane 1, Station 2.4

Modular Vane 2, Station 2.4

3-hole probe survey data, Station 2.2

100% Nc

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0.6 0.7 0.8 0.9 1 1.1 1.2

Fraction of span from hub

Local pressure coefficient, p0,2.2(z)/p0,2.2_average

85% Nc

90% Nc

95% Nc

100% Nc
NASA/TM—2015-218455 15

Figure 12.—Comparison of measured flow angle from 3-hole probe at Station 2.2 

and pre-test CFD at design point (100 percent Nc, ṁc,ex = 3 lbm/s).

Figure 13.—Comparison of impeller and stage adiabatic efficiency (ηTT) at 100 percent Ncas 

a function of inlet corrected flow from measurements at suppressed inlet conditions and 

pre-test CFD computed at standard day conditions (i.e., not corrected for suppressed inlet).

5.3.2 Impeller-Alone Efficiency

Impeller and overall stage efficiency at 100 percent Nc are provided in Figure 13. The measured 

impeller total-to-total adiabatic efficiency is near 87.5 percent throughout the flow range. The impact of 

diffuser/EGV losses is reflected in the substantial difference (4.5 to 10 points) between the impeller-alone 

and stage adiabatic efficiencies. Pretest CFD performance levels are provided in Figure 13 for context. 

Although the pre-test CFD shown was not corrected to account for suppressed inlet operation—refer to 

Table 5 for expected impact—a considerable shortfall in performance is evident.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

50 60 70 80 90 100

Fraction of span from hub

Swirl angle α2.2, deg.

Probe survey data, RR = 1.04

Design intent profile

100% Nc

60

65

70

75

80

85

90

95

10 10.5 11 11.5 12 12.5

Impeller and stage efficiency, %

Inlet corrected flow, lbm/s

Impeller (Sta. 0 to 2.4)

Stage (Sta. 0 to 3)

Impeller - pre-test CFD

Stage - pre-test CFD

100% Nc
NASA/TM—2015-218455 16

5.4 Diffuser, 90° Bend, EGV—Survey Data, Vane Loading, and Performance

5.4.1 Diffuser Exit Survey Data

Contours of measured total-pressure and flow angle from the three-hole probe surveys at Station 2.7 are 

provided in Figure 14 for design point operation conditions (Nc = 100 percent, ṁc,ex= 3 lbm/s). The surveys 

resolve one main-to-main passage, although two passages are reproduced in Figure 14 for visualization. The 

approximate positions of the main and splitter vanes are projected onto the contours for reference. Spanwise 

and pitchwise distortion is evident: the hub region is deficient in total pressure as is the suction side of the 

main blade (Passage A). A region of high swirl is evident on the shroud (to 20 percent immersion) on the 

suction side of the main blade, potentially indicative of separated or reversed flow. 

The impact of throttle setting on the total-pressure field at diffuser exit is illustrated in Figure 15, 

where the 100 percent Nc operating conditions at choke, design-point, and near-stall are presented. There 

is an evident redistribution of the total-pressure with throttle setting, with peak total-pressure rebalancing 

between Passages A and B. The corresponding swirl levels discharged from the diffuser section change 

little with throttle setting, though swirl does increase near the suction side (SS) trailing edge of the main 

vane as stall is approached. 

The transport of total pressure through the diffuser/90° bend, and the EGV was visualized using 

survey (Station 2.7) and rake data (Stations 2.8 and 3.0) in the diffusion system. Example results are 

provided in Figure 16 for design-point operation (100 percent Nc and ṁc,ex = 3 lbm/s). As evident by 

comparing the bottom two p0-distributions (Stations 2.7 and 2.8), the low-momentum flow exiting the 

diffuser near the hub is resolved by EGV rakes at lowest radius. It is also noted that it appears that much 

of the higher total-pressure flow located near the shroud region is not resolved by the EGV rake. 

Consequentially, 1-D annulus averaged values for p0 at the EGV inlet (not reported herein) may be 

underestimated by the rake layout. Better measurement resolution would be needed to quantify the 

uncertainty of this 1-D annulus average at the EGV leading edge.

Figure 14.—Contours of measured total pressure and flow angle at diffuser exit 

(Station 2.7) for design-point operation (Nc = 100 percent, ṁc,ex = 3 lbm/s).
NASA/TM—2015-218455 17

Figure 15.—Contours of measured p0, and flow angle α, at diffuser exit (Sta. 2.7) along design-point speed. (Nc = 100

percent) speedline during choke, design-point, and near-stall operation. (Note: legends are on different scales for p0.)

Figure 16.—Contours of measured p0 (Stations 2.7, 2.8, and, 3.0) for 

design-point operation (Nc = 100 percent, ṁc,ex = 3 lbm/s) through 

diffusion system from diffuser exit, to 90° bend, to EGV.

The total-pressure distribution at EGV exit (Station 3.0, Figure 16) is highly nonuniform (pitchwise) 

and is reflective of the high turning in the EGV and the position of the EGVs relative to the diffuser. 

Although not quantified, the predicted EGV exit flow field is provided as a pretest CFD line contour inset 

in Figure 16, which also reflects a high level of pitchwise segregation of total pressure: the rating 

indicates that one EGV passage is particularly weak. This flow deficiency is traceable to the wake of a 

main diffuser vane as noted during aerodynamic design (Ref. 2).

5.4.2 Diffuser Vane Loading

The loading diagrams for the diffuser main- and splitter-vanes are provided in Figure 17 for neardesign-point operation at inlet corrected flow of ṁc,in = 11 lbm/s. The experimental values are from static 

taps located around the main and splitter vanes (approximately 0.1 in. from vane surfaces) at the shroud 
NASA/TM—2015-218455 18

endwall; although not reported here, corresponding hub measurements were available as well. Evidently

the measured overall pressure rise at the near-design inlet corrected flow was lower than predicted. Note 

that the negative loading on the splitter vane reflects operation at negative incidence, in strong contrast to 

the lightly loaded splitter of pre-test CFD. The minimum pressures measured on the splitter pressure-side 

(near r = 10 in.) reflect higher velocities in Passage A than intended by design. In test, the positive 

loading on the main vane increases in the region of negative loading on the splitter (9.9 < r (in.) < 11.1).

(a)

(b)

Figure 17.—Comparison of predicted and measured loading diagrams (normalized static 

pressure versus radius) at the shroud on (a) diffuser main vane and (b) splitter vane, at near 

design-point operation (11 lbm/s, 100 percent Nc).

1.5

2

2.5

3

3.5

4

4.5

5

8.5 9 9.5 10 10.5 11 11.5

Normalized static pressure, 

p/p0,0

Radius, in.

Pressure side - pre-test CFD

Suction side - pre-test CFD

Pressure side - data

Suction side - data 100% Nc

1.5

2

2.5

3

3.5

4

4.5

5

8.5 9 9.5 10 10.5 11 11.5

Normalized static pressure, 

p/p0,0

Radius, in.

Pressure side - pre-test CFD

Suction side - pre-test CFD

Pressure side - data

Suction side - data
NASA/TM—2015-218455 19

5.4.3 Performance

The loss and pressure rise coefficients for the diffuser, the 90° bend plus EGV, and the overall 

diffusion system (diffuser, 90° bend, EGV) are provided in Figure 18 and Figure 19. The loss and 

pressure rise coefficients, given area-averaged 1-D inlet (“in”) and exit (“ex”) values, are obtained using 

ω = (p0,in − p0,ex )/(p0,in − pin ) and c p = (pex − pin )/(p0,in − pin ), respectively. The diffuser loss levels 

(Figure 18) are reduced as the stage is throttled to stall. Note that minimum diffuser loss occurs at a 

slightly lower exit-corrected flow than the ṁc,ex = 3 lbm/s design-intent. The loss coefficient in the 90°

bend plus EGV is relatively constant through choke and throughout much of the speed range. The 

minimum loss of each subcomponent of the stage does not occur at the same flow rate, partly contributing 

to the inability of the stage to achieve design-intent performance (see Figure 9 and Figure 10).

(a)

(b)

Figure 18.—Total-pressure loss coefficient across diffuser, 90° bend + EGV, diffuser + 

bend + EGV, as a function of (a) stage inlet and (b) stage exit corrected flow.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

10 10.5 11 11.5 12

Loss coefficient

Inlet corrected flow, lbm/s

Diffuser

90-deg. bend + EGV

Diffuser + 90-deg. bend + EGV

Stall Design Intent

(ṁc,ex = 3 lbm/s)

Choke

100% Nc

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

2.5 3 3.5 4 4.5 5

Loss coefficient

Exit corrected flow, lbm/s

Diffuser

90-deg. bend + EGV

Diffuser + 90-deg. bend + EGV

Stall Design

Choke

100% Nc
NASA/TM—2015-218455 20

The static pressure rise coefficients of the diffuser and the 90° bend plus EGV system are provided in 

Figure 19 as a function of inlet- and exit-corrected flow. The static pressure rise coefficient for the 90°

bend plus EGV is essentially constant between unchoke and stall.

(a)

(b)

Figure 19.—Static pressure rise coefficient across diffuser, 90° bend + EGV, diffuser + 

bend + EGV, as a function of (a) stage inlet and (b) stage exit corrected flow.

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

10 10.5 11 11.5 12

Static pressure rise coefficient

Inlet corrected flow, lbm/s

Diffuser

90-deg. bend + EGV

Diffuser + 90-deg. bend + EGV

Stall

Design-intent 

(ṁc,ex = 3 lbm/s)

Choke

100% Nc

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

2.5 3 3.5 4 4.5 5

Static pressure rise coefficient

Exit corrected flow, lbm/s

Diffuser

90-deg. bend

EGV

100% Nc

Choke

Stall Design
NASA/TM—2015-218455 21

6.0 Compressor Sensitivities

6.1 Reynolds Number Sensitivity

As presented in Section 5.0, facility limitations required testing at reduced plenum pressures. 

Operation at reduced pressure and, hence, Reynolds number, impacted performance detrimentally as 

noted in Table 5. A Reynolds number, Re sensitivity was conducted at 85 percent Nc by varying plenum 

pressure between 10 ≤ p0,0 ≤ 13 psia, effecting a 25 percent change in Re. Once the plenum pressure was 

established, a typical speed line from choke to near stall was run. The effect of Reynolds number on 

measured stage adiabatic efficiency was evaluated against a power-law correlation of efficiency loss to 

Reynolds number relationship typically used in compressor testing.

n

tt ref ref

tt

−





 



 = − η

− η

Re

Re

1

1

,

The exponent, n, represents the rate of efficiency loss with Reynolds number for a specific compressor. 

The reference condition in our case was for a plenum pressure of 13.0, and the compressor operating at an 

exit corrected flow condition of 3.0 lbm/s. A curve fit of the three tested data points results in n = 0.077 

for this machine, typical of previous findings (Ref. 8) and turbulent flow scaling (n = 0.1). It is noted that 

the pre-test CFD in Table 5 predicted a greater sensitivity (n  0.3).

6.2 Tip-Clearance Sensitivity

Compressor sensitivities to tip-clearance derivatives of flow rates, stage pressure ratio, stage 

efficiency and overall stability margin were obtained. Tip gaps of ε = 0.012, 0.015 and 0.018 in.

corresponding to reduced clearances, ε/b, of 2, 2.5 and 3 percent for this impeller with a trailing edge span 

of b = 0.609 in. Speed lines from choke to stall were run at 100 percent Nc speed, with the stall point 

being identified as the last stable operating point that could be obtained. Performance sensitivities were 

assessed at the exit corrected flow rate of 3.0 lbm/s, and provided in Table 6, with the detailed speed lines 

documented in Figure 20(a) and (b). 

The reduction in efficiency of 0.24 points for a 1 percent change in ε/b is consistent with findings of 

Skoch et al. (Ref. 8), but the increase in surge margin as the clearance increases is not a typical result. As 

noted in Reference 3, this SM response with clearance may be reflective of hub endwall weakness in the 

stability limiting element. For a 1 percent change in ε/b, the pressure rise and inlet corrected mass flow 

rates respond with reductions of 0.44 and 0.47 points, respectively. 

TABLE 6.—TIP CLEARANCE SENSITIVITIES FROM

LINER CURVE FIT OF ∆y/y RELATIVE TO ∆ɛ/b

Tip clearance,

ɛ/b

Surge margin Adiabatic 

efficiency

p0,3/p0,0 Inlet 

corrected 

flow

0.02 7.5% 0.822 4.69 10.85

0.025 7.8% 0.821 4.67 10.80

0.03 7.9% 0.819 4.65 10.77

Sensitivities –0.24 –0.44 –0.47
NASA/TM—2015-218455 22

(a)

(b)

Figure 20.—Tip-clearance sensitivity (a) pressure ratio, and (b) adiabatic efficiency.

3.8

4.0

4.2

4.4

4.6

4.8

5.0

10.0 10.5 11.0 11.5 12.0

Pressure Ratio, 

p0,3.0/p0,0

Inlet Corrected Flow, lbm/s

Design, ɛ = 0.012", ɛ/b=2.0

ɛ = 0.015", ɛ/b=2.5

ɛ = 0.018", ɛ/b=3.0

3.0 lbm/s Operating Line

0.72

0.74

0.76

0.78

0.80

0.82

0.84

10.0 10.5 11.0 11.5 12.0

Adiabatic Efficiency, ηtt

Inlet Corrected Flow, lbm/s

Design, ɛ = 0.012", ɛ/b=2.0

ɛ = 0.015", ɛ/b=2.5

ɛ = 0.018", ɛ/b=3.0

3.0 lbm/s Operating Line
NASA/TM—2015-218455 23

7.0 Discussion

As noted previously, the design-intent performance of the HECC stage was not achieved. The 

performance shortfall is summarized well in Table 5 where the adiabatic total-to-static efficiency of the 

stage is shown to be 3.46 points lower than design-intent, pre-test CFD corrected for suppressed inlet. The 

miss in achieved stage static pressure rise reflects low impeller efficiency (Figure 13) and diffuser/EGV 

performance as evidenced by the high stage (EGV) exit Mach numbers and (inferred) swirl angles 

(Table 5). The stability margin of the test article was 7.5 percent rather than 12 percent design-intent and 

the ṁc,in flow rate was approximately 2.25 percent lower (see Figure 13) than design-intent predictions.

Though CFD and data comparisons are beyond the scope of the present paper, it is worth comparing 

the predicted and measured diffuser loss buckets shown in Figure 21. The shift in corrected flow between 

the predicted and measured diffuser loss buckets indicates that the test diffuser does not achieve its 

minimum loss at the design intent flow rate, as further explored in Figure 22.

As illustrated in Figure 22(a), the predicted and experimental corrected-flow characteristic for the 

diffuser are in fairly good agreement, and the match at the design-intent exit corrected flow (3 lbm/s) is 

excellent; that is, when the diffuser/EGV is considered in isolation, it processes flow it is supplied as per 

design intent. The impeller corrected flow characteristics (CFD vs. experiment) are provided in 

Figure 22(b). Unlike the diffuser/EGV system, there is a clear difference between the predicted and 

measured impeller inlet and exit corrected flows; this is reflective of the reduced measured efficiency of 

the impeller (Figure 13), leading to higher impeller exit (diffuser inlet) corrected flows than intended. This 

error makes the diffuser appear to be undersized (insufficient flow area). 

Figure 21.—Normalized pressure drop across diffuser (test vs. design-intent), showing 

mismatch in physical flow at minimum in loss bucket.

0

5

10

15

20

25

30

35

10 10.5 11 11.5 12 12.5

∆

p0/p0 of diffuser, %

Inlet corrected flow, lbm/s

NASA HECC 12 mil - coarse grid, mixing-plane

NASA HECC 12 mil - fine grid - mixing-plane

HECC data

Design-intent 

(ṁc,ex = 3 lbm/s)

100% Nc

Expt.

stall boundary
NASA/TM—2015-218455 24

(a)

(b)

Figure 22.—(a) Diffuser corrected flows and (b) Impeller corrected flow, showing a mismatch 

in impeller and diffuser corrected flow capacity.

2

2.2

2.4

2.6

2.8

3

3.2

2 2.5 3 3.5 4

Impeller exit corrected flow, lbm/s

Stage exit corrected, lbm/s

NASA HECC 12 mil - coarse grid - mixing-plane

NASA HECC 12 mil - fine grid - mixing plane

HECC test data

Design-intent

100% Nc

2

2.2

2.4

2.6

2.8

3

3.2

10 10.5 11 11.5 12 12.5

Impeller exit corrected flow, lbm/s

Stage (impeller) inlet corrected flow, lbm/s

NASA HECC 12 mil - coarse grid - mixing-plane

NASA HECC 12 mil - fine grid - mixing-plane

HECC test data

Design-intent 

(ṁc,ex = 3 lbm/s) 100% Nc
NASA/TM—2015-218455 25

8.0 Summary

The aerodynamic performance of a new centrifugal compressor stage, designed to meet a high 

performance goal set consistent with future turboshaft engine needs, was established by experimental 

testing. The experimental results presented are associated with an openly available geometry (Ref. 3). In 

addition to providing future insight into technical challenges associated with compact, highly loaded 

centrifugal compressors, the dataset is intended to contribute value to the turbomachinery community in 

terms of computational/modeling tools assessment and development.

The stage and subcomponent performance levels were described. The achieved performance and 

stability was found to be lower than design-intent. Detailed root-cause-analysis of the performance 

shortfall was initiated within the NRA contract (Ref. 3) under which the work was conducted. These 

analyses continue, with intent to inform future design efforts.

References

1. Welch, G. E., Hathaway, M. D., Skoch, G. J., and Snyder, C. A., “Rotary-Wing Relevant Compressor 

Aero Research and Technology Development Activities at Glenn Research Center,” Proc. AHS Int. 

65th Annual Forum, May, 2009; also NASA/TM—2012-217280, Mar. 2012.

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

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

Applications,” AHS Int. 67th Annual Forum, May 2011.

3. Medic, G., Sharma, O., Jongwook, J., Hardin, L. W., McCormick, D. C., Cousins, W. T., Lurie, E. A., 

Shabbir, A., Holley, B. M., and Van Slooten, P. R., “High Efficiency Centrifugal Compressor for 

Rotorcraft Applications,” NASA/CR—2014-218114, Nov. 2014.

4. Brokopp, R. A. and Gronski, R. S., “Small Engine Components Test Facility Compressor Testing 

Cell,” AIAA-92-3980, July, 1992; also NASA TM-105685, 1992.

5. McCain, T. F., and Holbrook, G. J., “Coordinates for a High Performance 4:1 Pressure Ratio 

Compressor,” NASA CR-204134, July 1997

6. ASME, “PTC 19.5-2004 Performance Test Codes,” 2005

7. Giel, P.W., Thurman, D. R., Lopez, I., Boyle, R. J., Van Fossen, G. J., Jett, T. A., Camperchioli, W.

P., and La, H., “Three Dimensional Flow Field Measurements in a Transonic Turbine Cascade,” 

ASME 96-GT-113, June 1992

8. Skoch, G. J., Moore, R. D., “Performance of Two 10-lb/sec Centrifugal Compressors With Different 

Blade and Shroud Thicknesses Operating Over a Range of Reynolds Numbers,” AIAA 87-1745, July 

1987.