Rotary-Wing Relevant Compressor Aero Research 

and Technology Development Activities 

at Glenn Research Center

NASA/TM—2012-217280

March 2012

Gerard E. Welch, Michael D. Hathaway, and Gary J. Skoch

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

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

ARL–TR–4757

U.S. ARMY

RESEARCH LABORATORY
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Rotary-Wing Relevant Compressor Aero Research 

and Technology Development Activities 

at Glenn Research Center

NASA/TM—2012-217280

March 2012

Gerard E. Welch, Michael D. Hathaway, and Gary J. Skoch

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

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

ARL–TR–4757

U.S. ARMY

RESEARCH LABORATORY

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

65th Annual Forum and Technology Display (AHS Forum 65)

sponsored by the American Helicopter Society

Grapevine, Texas, May 27–29, 2009
Acknowledgments

This effort is sponsored by the Subsonic Rotary Wing project of the NASA Fundamental Aeronautics Program. Portions of this 

work were conducted under the ARL Army Stall Control Engine Demonstrator program, funded through the Army Research Office.

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Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

either expressed or implied, by the National Aeronautics and 

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This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217280 1

Rotary-Wing Relevant Compressor Aero Research and Technology 

Development Activities at Glenn Research Center

Gerard E. Welch, Michael D. Hathaway, and Gary J. Skoch

U. S. Army Research Laboratory

Glenn Research Center

Cleveland, Ohio 44135

Christopher A. Snyder

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Technical challenges of compressors for future rotorcraft 

engines are driven by engine-level and component-level 

requirements. Cycle analyses are used to highlight the enginelevel challenges for 3000, 7500, and 12000 SHP-class engines, 

which include retention of performance and stability margin at 

low corrected flows, and matching compressor type, axial-flow 

or centrifugal, to the low corrected flows and high temperatures 

in the aft stages. At the component level: power-to-weight and 

efficiency requirements impel designs with lower inherent 

aerodynamic stability margin; and, optimum engine overall 

pressure ratios lead to small blade heights and the associated 

challenges of scale, particularly increased clearance-to-span 

ratios. The technical challenges associated with the aerodynamics of low corrected flows and stability management impel the 

compressor aero research and development efforts reviewed 

herein. These activities include development of simple models 

for clearance sensitivities to improve cycle calculations, fullannulus, unsteady Navier-Stokes simulations used to elucidate 

stall, its inception, and the physics of stall control by discrete 

tip-injection, development of an actuator-duct-based model for 

rapid simulation of nonaxisymmetric flow fields (e.g., due inlet 

circumferential distortion), advanced centrifugal compressor 

stage development and experimentation, and application of stall 

control in a T700 engine.

Nomenclature

AATE Advanced Affordable Turbine Engine

b impeller exit blade span

c chord

h blade height (span)

g staggered gap

JHL Joint Heavy Lift

LCTR Large Civil Tilt-Rotor

m , mc  mass flow rate, corrected mass flow rate

OPR overall pressure ratio

SHP shaft horsepower

SFC specific fuel consumption

SP specific power

T3 compressor discharge temperature

T4 turbine entry temperature

δ* aerodynamic blockage

ε clearance (e.g., in mils)

η efficiency

ψ pressure rise coefficient

φ flow coefficient

Introduction

An overview of rotorcraft-relevant compressor research and 

technology development activities underway at Glenn 

Research Center is provided herein. The research activities are 

sponsored by the NASA Subsonic Rotary Wing (SRW) 

project and are intended to address key aerodynamic 

challenges associated with the compressors of the type used in 

advanced turboshaft engines of the 3000 (DoD AATE

(Ref. 1)), 7500 (NASA LCTR (Refs. 2 and 3)), and 12000 

(DoD JHL, cf. Ref. 4) SHP-class. A particular objective of the 

described research is to support technology development of 

the 7500 SHP-class engines for the NASA LCTR mission.

The principal challenge posed by the LCTR mission is the 

requirement to reduce the main rotor speed from 650 ft/s at 

ground take-off to 350 ft/s at Mach 0.5 cruise (Acree et al.

(Ref. 3)) Although the rotor speed is reduced by approximately 50 percent, the specific power levels (SHP/lbm/s) at the 

take-off and cruise points are essentially the same. This 

requirement for wide variability in rotor speed represents a 

significant departure from the nearly constant-speed operation 

of conventional rotorcraft. The required variability impels 

research and technology efforts related to (i) variable-speed 

transmission and/or (ii) variable-speed power-turbine output.

While the power-turbine shaft speed can vary by a factor of 

two in the latter approach, preliminary cycle analyses of the 

LCTR mission indicate that the gas generator components 

(compressor and HPT) operate at approximately constant 

corrected conditions and no additional operability requirements are imposed on the compression system (Snyder and 

Thurman (Ref. 5)).

The technical challenges for the compressors are related to 

aerodynamics of low corrected flows ( m c ∝ m T p ) in aft 
NASA/TM—2012-217280 2

stages, strength-of-materials and cooling limitations at 

elevated compressor discharge temperatures, and stability.

These challenges are discussed in the first part of the paper. 

This is followed by an overview of on-going modeling, 

component testing, and engine testing activities intended to 

address these challenges.

Compressor Technology Challenges

The technical challenges for the compressors are driven by

engine system and component level requirements.

Engine-Level Challenges

At the engine level, the high power-to-weight and fuel 

efficiency requirements of rotorcraft engines push overall 

pressure ratios (OPRs) to increasingly higher levels. Example 

curves for SFC and specific power (SP) as functions of OPR for 

the 3000, 7500, and 12000 SHP engine classes are shown in 

Figure 1. These results were obtained using the Numerical 

Propulsion System Simulation (NPSS) code (Jones (Ref. 6)), for 

the fixed turbine entry temperatures indicated in Figure 1,

specified technology correlations for turbine cooling effectiveness as a function of cooling flow (Gauntner (Ref. 7)), and 

compressor polytropic efficiency as a function of compressor 

exit-corrected flow (Figure 2). The influence of the higher 

temperature cooling air and degradation of compressor 

efficiency in the aft-stages at increased OPR, leads to the optima 

in OPR, in terms of SFC minimization, as shown in Figure 1(a).

(a) SFC versus OPR (b) SP versus OPR

Figure 1.—Engine cycle calculation results, showing (a) SFC and (b) specific power as a function of OPR for the 3000, 7500, and 

12000 SHP-class engines.

(a) Technology curve basis (b) Technology curves

Figure 2.—Notional compressor technology curves showing (a) design-point polytropic efficiency as function of exit-corrected flow 

for centrifugal and axial-flow compressor stages, and representative aggregate performance curve (green/dashed) used in cycle 

analyses herein; and, (b) projected improvements in notional aggregate compressor efficiency levels to be achieved by aerodynamic improvements.

0.0

50.0

100.0

150.0

200.0

250.0

300.0

350.0

0 10 20 30 40 50 60

Overall pressure ratio

Specific power, SHP/lbm/s

3000 SHP, T4 = 2700 F

7500 SHP, T4 = 3000 F

12000 SHP, T4 = 3000 F

 

0.82

0.84

0.86

0.88

0.9

0.92

0.94

0 1 2 3 4 5

Exit-corrected flow, lbm/s

Polytropic efficiency

Exit corrected flow

Centrifugal 

compressors

Axial compressor

Notional aggregate

compressor technology curve

ηp

 

Improved aero performance Future

Fielded
NASA/TM—2012-217280 3

Figure 3.—Exit-corrected flow in example 3000, 7500, and 12000 SHP engines as a function of OPR.

Figure 4.—Compressor discharge temperature (T3) in example 

3000, 7500, and 12000 SHP engines as a function of OPR, 

showing opening of OPR-space accessible by higher 

temperature materials capabilities.

The high engine OPRs introduce technical challenges 

associated with the aerodynamics of low corrected flows and 

blade heights (Figure 3(a) and (b)) and high compressor 

discharge temperatures (T3, Figure 4). In engines of a fixed 

SHP level, these challenges only worsen as T4 (and hence SP) 

is increased, leading to engines of lower mass flow rate and 

higher optimum OPR. The challenges associated with low 

corrected flows and high T3 are highlighted in the following 

sections, engine by engine.

3000 SHP-Class Engines

At the 3000 SHP (AATE) level with physical flow rates of 

11 to 13 lbm/s, aft stage exit-corrected flow rates ( mc,3  ) can be 

less than 0.75 lbm/s (see Figure 3(a)). At these flow sizes and 

associated small blading (Figure 3(b)), the centrifugal 

compressor overtakes the axial compressor in terms of 

efficiency, as shown notionally in Figure 2(a), compactness, 

and weight. The cross-over point in terms of exit-corrected 

flow (e.g., 1.5 lbm/s), while dependent weakly on aero 

technology levels and engine size, is the point at which the 

aerodynamics of the axial stages is degrading rapidly—due to 

the effects of relatively large clearance-to-span, bladetolerances, and relative boundary-layer (aero-blockage) 

fractions. An important technical challenge in the 3000 SHP 

engine size is related, in part, to retaining high efficiency in 

compact, highly loaded axial and centrifugal compressors with 

low corrected flows. Overcoming this challenge impels 

research related to management of loss associated with large 

impeller-tip/shroud clearance-to-exit-span ratios, relative 

fillets, and blade thickness (cf. scaling study of Skoch and 

Moore (Ref. 8)). Considering the stress levels sustained by aft 

stage centrifugals, engine pressure ratio—or specifically the 

associated compressor discharge temperatures, T3—may well 

be limited by the strength of impeller materials required to 

meet life (as suggested in Figure 4).

7500 SHP-Class Engines

In the 7500 SHP-class engines, the physical weight flows 

(e.g., 25 lbm/s) and optimum OPR (e.g., 37:1) lead to 

compressor exit-corrected flows near unity (see Figure 3(a)). 

This power class engine may prove particularly challenging to 

compressor aero. At this corrected flow, centrifugal stages 

would be preferable to axial stages (Figure 2(a)); however, the 

T3s associated with these OPRs might prove incompatible with 

the high stress levels of centrifugal impellers, and axial stages 

might need be used, even at the relatively low aft-stage 

corrected flows (< 1 lbm/s). The technical challenges in the aft 

axial stages are thus related to maintaining high efficiency 

a. Exit-corrected flow vs. OPR b. Exit blade height vs. OPR

0

0.5

1

1.5

2

2.5

3

3.5

0 10 20 30 40 50 60

Overall pressure ratio

Exit blade height, inches

3000 SHP

7500 SHP

12000 SHP

0

0.5

1

1.5

2

2.5

3

3.5

4

0 10 20 30 40 50 60

Overall pressure ratio

Exit-corrected flow, lbm/s

3000 SHP

7500 SHP

12000 SHP

0

200

400

600

800

1000

1200

1400

1600

0 10 20 30 40 50 60

Overall pressure ratio

Compressor discharge temperature, 

3000 SHP

7500 SHP

12000 SHP

Current T3

limit (notional) 

for centrifugal 

compressors

T3 limit with 

higher strength materials

Compressor discharge temperature (T3), °F
NASA/TM—2012-217280 4

with small blade heights (< 0.5 in.). This technical challenge is 

a subject of research in a companion NASA project (Subsonic 

Fixed Wing) related to the aerodynamics of low-corrected 

flows in aft axial stages of high OPR ultra-high bypass ratio 

engines. The blade heights of these engines are envisaged 

comparable to those shown for the 7500 SHP-class engine 

(Figure 3(b)). As suggested by Figure 4, if/when impeller/backplate materials with increased strength at temperature 

are employed, use of a centrifugal compressor to replace aft 

axial stages may be warranted in this engine class.

12000 SHP-Class Engines 

The 12000 SHP-class engines (JHL) will have relatively

high physical mass flow rates (e.g., 45 lbm/s) at high OPRs 

(e.g., 42:1) leading to exit-corrected flows of about 1.5 lbm/s. 

At these corrected flows and high temperatures, an all-axial 

compressor would be expected. The technical challenges are 

again related to the aerodynamics of low-corrected flow in 

axial aft-stages.

Component Level Challenges

The impact of compressor polytropic efficiency on LCTRclass (7500 SHP) engine efficiency (fuel consumption), size 

(weight), and available turbine cooling temperatures (T3) is 

shown in Figure 5. Considering an example OPR of 37:1 (see 

Figure 1(a)), a two point increase in polytropic efficiency,

from “fielded” to “future” technology levels (Figure 2(b) viz. 

Figure 3(a)), leads to a 3 percent reduction in fuel consumption (Figure 5(a)) and a 5 percent reduction in engine mass 

flow rate (Figure 5(a)). Additionally, the same technology 

improvement would lead to 35 °F cooler compressor discharge 

temperature at the 37:1 OPR (Figure 5(b)). These benefits

highlight the importance of addressing technical challenges 

associated with improving component efficiency by managing 

blade row losses (3-D aero designs), clearances, and leakages, 

and insuring stage matching so that blade rows operate as per 

design-intent.

Power-to-weight requirements force high aerodynamic 

loading levels associated with compactness (low weight) 

requirements. Concurrently, the aero-loading levels must be 

tempered by the component efficiency and stall margin 

required to meet engine efficiency and operability. These 

factors push designers to stage designs with lower inherent 

aerodynamic stability margin—a design trade which may be 

acceptable in advanced engines if active stability management 

methodologies are realizable (cf. Larosiliere et al. (Ref. 9)). 

The technical challenges then associated with highly efficient, 

high power density compressor components involve development of efficient and reliable stability management, or stall 

control, technologies.

As the corrected flows decrease with increasing OPR 

(Figure 3(a)) and the aft-stage blade heights are reduced 

(Figure 3(b)): The aft axial stages are pushed to higher hub-totip ratios in general, the blade counts increase, and/or the aspect 

ratios decrease; the minimum blade thicknesses and fillets at 

the small sizes and high temperatures may well be set by 

material/strength requirements or machining tolerances rather 

than aero considerations; and, the impact of endwall and 

clearance flows takes on relatively stronger importance. The 

management of endwall aerodynamic blockage levels, in 

particular, becomes increasingly challenging as the relative 

clearances (ε/h) increase.

Figure 5.—Impact of two-point increase in compressor polytropic efficiency (cf. Figure 2(b)) from fielded to future engine in terms 

of (a) ratios of specific fuel consumption and mass flow rate; and (b) compressor discharge temperature, T3.

0.91

0.92

0.93

0.94

0.95

0.96

0.97

0.98

0.99

1

0 10 20 30 40 50 60

Overall pressure ratio

Ratio of future to fielded

SFC ratio

Mass flow rate ratio

-50.0

-45.0

-40.0

-35.0

-30.0

-25.0

-20.0

-15.0

-10.0

-5.0

0.0

0 10 20 30 40 50 60

Overall pressure ratio

Reduction in T3, F

a. SFC and mass flow rate ratios b. Reduction in compressor discharge temperature Reduction in T3, °F
NASA/TM—2012-217280 5

Compressor Research Activities

Specific computational and experimental research activities 

described below are intended to address the technical 

challenges associated with the aerodynamics of aft stages with 

low corrected flows and compressor stability. With regards to 

compatibility with low corrected flows, a rudimentary 

clearance flow modeling effort is first described, followed by 

a description of centrifugal compressor research activities in 

the NASA Small Engine Components Test Facility. With 

regards to stability, three activities are then described: T700 

engine/stall-control testing and supporting unsteady NavierStokes computations; and, the development of a new actuatorduct based nonaxisymmetric throughflow model.

Modeling for Improved Engine Cycle Calculations

The reduction in aerodynamic performance at low corrected 

flows must be accounted for in the engine performance 

calculations (as was approximated herein through the 

compressor technology curve, Figure 2(b)). The impact of 

rotor/impeller tip clearance (ε/h or ε/b), in particular, is critical 

given its impact on both compressor performance and 

stability. Typical clearance sensitivities for pressure ratio, 

efficiency, flow, flow range, and stability are documented in 

the literature, and are shown qualitatively in Figure 6(a). As 

the clearance-to-chord (or span, ε/h or ε/b) increases, stage 

efficiency, stability margin, pressure ratio, and flow range 

generally decrease. Cumpsty (Ref. 10) has summarized key 

references for clearance sensitivities for both centrifugal and 

axial compressors. For example, Wisler (Ref. 11) has reported 

for a low-speed four-stage axial compressor that an increase in 

clearance-to-chord from 1.6 to 3.4 percent led to a 1.5 percent

reduction in efficiency, and 11 percent reduction in flow 

range, and a 9.7 percent reduction in peak pressure rise, while 

Freeman (Ref. 12) found a 1.4 percent decrease in efficiency 

for a one point change in clearance to chord in a high-speed 

multistage compressor. Similarly, in centrifugal compressors, 

roughly 0.25 percent in efficiency is sacrificed for every 1

percent clearance to exducer-span (ε/b, cf. Skoch and Moore

(Ref. 8)). In terms of engine cycle modeling, the clearance 

sensitivities are of interest for both design-point performance 

estimates and for transients (i.e., case and rotor differential 

cooling effects). In terms of design-point estimates, as OPR 

increases the ratio of an acceptable running clearance (e.g., 

15 mils) to the aft-stage blade height (see Figure 3(b)) 

increases strongly. For example, if the ratio of a clearance to

chord of an LCTR (7500 SHP) engine at 20:1 (and m c,ex ≈

1.7 lbm/s) was ε/c ≈ 2 percent, then using Figure 3(b) the same 

aspect ratio blading at 40:1 ( m c,ex ≈ 1) would have 

ε/c ≈ 3 percent. In light of the sensitivities cited for the axial

compressors above, the loss in aft-stage efficiency attributable 

to reduced relative clearances alone for this increase in OPR 

would be expected to be 0.8 to 1 point (cf. Figure 2(b)).

A simplified model to account for the impact of increased 

relative clearances is under development using documented 

sensitivities of this kind. The compressor characteristic (ψ – φ

and η – φ) at a given corrected speed are determined 

(see Robbins and Dugan (Ref. 13)) from known (or designintent) performance characteristics and clearance (ε/c). 

Herein, the pressure rise characteristic is modeled as

Figure 6.—Impact of increased clearances on compressor performance, showing (a) expected impact on pressure rise, 

efficiency, stall margin, flow, and flow range, and (b) modeled impact of clearance on performance curves intended 

for use in cycle calculations.

op lines

stability

boundaries

PR

η

mc,3 

1

1.1

1.2

1.3

1.4

1.5

1.6

1.7

0 0.1 0.2 0.3 0.4 0.5 0.6

Corrected flow function

Pressure ratio

1%

2%

3%

a. Impact of clearance at constant speed b. Model results

Clearance-to-chord

50% speed

Tight clearances 100% speed

Open clearances
NASA/TM—2012-217280 6

2

min

max 1 1













 − φ

φ ψ ψ = − R ⋅ , where ψmax is the pressure 

coefficient at stall (maximum), φmin is the stalling flow 

coefficient, and max

2

min

2

2

2

1

ψ

φ⋅ ∂φ

∂ ψ R = − sets the curvature of the 

characteristic. For a given clearance, the stalling pressure ratio 

as a function of percent ε/c is determined from the correlation 

of Smith (Ref. 14), which documents a 4.6 percent peak 

pressure loss for each 1 percent clearance-to-chord. The 

aerodynamic blockage (δ*

) at this stability boundary is 

estimated using the correlation of Koch and Smith (Ref. 15) 

and is a function of the clearance to stagger-gap ratio (ε/g). 

The percent of span corresponding to the aerodynamic 

blockage fraction (δ*

/h) is assumed to do no work and the 

stalling flow rate is back calculated using the known 

compressor characteristic and the stalling pressure ratio from 

the Smith correlation. 

The model is in its inception and is certain to undergo 

refinement during its implementation and validation. 

Preliminary results in Figure 6(b) indicate the calculated 

impact of clearance variations on the performance characteristics at 50 and 100 percent corrected speeds. The clearance 

strongly reduces the stability margin and maximum pressure 

rise capability of the compressor. The efficiency is reduced 

concomitantly, according to the sensitivities described above 

(e.g., 1.4 percent decrease in η for each 1 percent increase in 

clearance-to-chord).

Experiments in the Small Engine Components 

Test Facility

Experimental research on both axial and centrifugal 

compressor aerodynamics has been conducted in the NASA 

Small Engine Components Test Facility (SECTF, see Brokopp 

and Gronski (Ref. 16)) over the past 25 years. The SECTF 

(Figure 7(a)) is ideally suited for testing rotorcraft-relevant 

single and multistage compressors. A 6000-hp variable 

frequency drive motor and gearbox is used to drive the 

research article at speeds up to 60,000 rpm. The throttle valve 

and exhaust sprayer cooler are rated for operation at pressure 

ratios up to 30:1. Inlet air pressure can be varied from 2 to 50 

psia and temperature from ambient down to –50 °F. Maximum 

flow capacity, dependent on inlet conditions, is in the range 40 

to 65 lbm/s. The rig, as configured currently, can accommodate 

compressors up to 20 in. diameter. 

Skoch and Moore carried out a centrifugal compressor 

scaling study during the late ‘80s which documented 

efficiency sensitivities to variations in clearance (ε/b), fillet 

size, and Reynolds number (Ref. 8). Laser Doppler velocimetry (LDV) was subsequently used to characterize the impeller 

discharge flow and vane-island diffuser flow in a series of 

experiments by Skoch et al. (Ref. 17) and Wernet et al.

(Ref. 18) The laser anemometry efforts provided valuable flow 

field data sets for code validation (see, for example, 

Larosiliere et al. (Ref. 19)), and were followed by a series of 

stall control experiments (discussed below). In addition to 

small (2.5 lbm/s) and large (10 lbm/s) centrifugal compressors, 

mc,in  = 10.5 lbm/s ( m c,ex ≈ 2.75 lbm/s), 2.5-stage with 5:1 

design-point pressure ratio was tested in the SECTF facility as 

well (Adamczyk et al. (Ref. 20).

The SECTF went into standby in 2003. Under the current 

NASA Subsonic Rotary Wing project, the test cell has been 

restored to an operational state. This investment is to enable 

in-house research of advanced centrifugal compressors with 

open geometry, to provide a facility for research formulated 

with industry and academia, and for DoD/industry collaboration. The SECTF is currently being used to re-baseline the 4:1 

Figure 7.—(a) photo of SECTF and (b) CC3 cross-section showing shroud stall control injection scheme.

a. Small Engine Component Test Facility b. Cross-section of CC3 compressor stage
NASA/TM—2012-217280 7

CC3 compressor stage (Figure 7(b)) comprising the impeller, 

vane-island diffuser, and 90° turning duct (no deswirl vanes) 

(Ref. 17). The 10 lbm/s CC3 compressor is a scaled-up version 

of a 3.6 lbm/s stage that facilitates detailed flow anemometry. 

Following baseline performance measurements, the 50 mil 

variable clearance capability of SECTF will be exercised to reestablish clearance sensitivities for performance and range. 

Current entry plans also include acquisition of r – θ surveys 

downstream of the vane-island diffuser—a first in CC3—and 

radial surveys of high-response total-pressure in the vaneless 

space between the impeller and diffuser.

NASA Research Announcement Contract on 

Advanced Centrifugal Compressors

A 2007 NASA Research Announcement (NRA) solicitation 

included requests for proposals to support advanced 

centrifugal compressor research and development with an aim

to improve design methodologies and component performance 

levels for future rotorcraft applications. The objectives of the 

3-year activity were to include the following:

• Identify and prioritize key knowledge gaps, and outline 

experimental testing needed to advance the state-of-theart of rotorcraft-relevant (Table 1) centrifugal compressor technology; 

• Determine research measurements required for tool 

validation and for needed insight into salient flow physics;

• Design, fabricate, install and collaboratively test a 

advanced, rotorcraft-relevant, centrifugal compressor 

research test article in the SECTF to obtain detailed aerodynamic and aeromechanical measurements; and,

• Acquire high-quality research measurements needed to 

clarify flow physics phenomena and to establish detailed 

data sets for development and validation of new methods.

A 3-year NRA contract was awarded to United Technologies 

Research Center (UTRC) in late 2008. A research centrifugal 

compressor stage—impeller, diffuser, turning duct, and deswirl 

vanes—which has key attributes (cf. Table 1) appropriate to 

state-of-the-art rotorcraft engines, is to be designed, built, and 

tested in the SECTF as a follow on to current CC3 experimentation. Table 1 reflects a highly aggressive compressor design in 

terms of the combined efficiency, work factor, stall margin, and 

compactness (weight) requirements. Like CC3 (Ref. 17), the 

new test article will be scaled-up to the 10 lbm/s flow size so as 

to accommodate detailed flow field measurements. The scaling 

will maintain tolerances, relative fillet sizes, and blade thickness 

to insure aerodynamic similitude with a 2.5 to 3 lbm/s rotorcraftrelevant application compressor. The centrifugal stage is 

scheduled for delivery and installation during FY10 and checkout testing and baselining at the end of FY11; thereafter, the test 

article is intended to be used by the government, industry, and 

academia to acquire key consensus data sets.

TABLE 1.—NRA SOLICITATION—APPLICATION 

COMPRESSOR DESIGN-POINT CHARACTERISTICS

Metric Value or Range

Stage pressure ratio 4.5 < PR < 6

Inlet corrected flow 2.5 ≤ m c,in ≤ 3 lbm/s

Stage-exit corrected flow 0.7 ≤ m c,ex ≤ 0.8 lbm/s

Work factor 0.60 ≤ ∆ ≤ 2 0 / H Utip 0.75

Polytropic stage efficiency ≥ 0.88

T3 capability 950 < T (°F) < 1000

Max flow path diameter to impeller tip diameter rmax /rtip ≤ 1.45

Design stability margin 25%

Stall Control Activities

The Active Stall Control Engine Demonstration (ASCED) 

project was initially funded by the U.S. Army Research 

Laboratory (ARL) and was completed with augmentation 

funding from the Subsonic Rotary Wing Project of the NASA 

Fundamental Aeronautics program. The project had both 

engine testing (experimental) and computational modeling 

aspects (next section). The objective of the project was to 

demonstrate stall control technologies, which had been 

developed earlier in NASA axial and centrifugal compressor 

component experiments, in an engine. The intent was to 

understand the impact of integration and operation in the 

engine environment on achieved flow range extension. In the 

following sections, the key findings from the earlier 

components experiments are first provided, followed by 

overviews of on-going experimental and computational stall 

control activities.

Earlier NASA Component Experiments

Suder et al. (Ref. 21) demonstrated flow range and stability 

improvements in transonic, single-stage, axial-flow compressors and fans using steady-state injection through discrete

injectors upstream of the compressor rotor. The discrete 

injectors were distributed circumferentially about the 

compressor case and were operated in choke. The emitted jets 

effectively unload the rotor tip regions and in this way reduce

the aerodynamic blockage levels at the case that lead to stall. 

Suder et al. (2001) (Ref. 21) determined that the arrangement 

of the injectors in terms of their relative circumferential 

position was less important to flow range extension than were 

the number of injectors and the axial velocity of each jet. A 

Coanda injector, suitable for integration between blade rows 

was demonstrated as well and later assessed in a highperformance multistage compressor (Strazisar et al. (Ref. 22).

Skoch extended the steady injection work just described to 

centrifugal compressors (Refs. 23 and 24) and demonstrated 

stability margin enhancement (+1.7 percent) using injectors 

that directed air-jets along the shroud surface in the vaneless 

region between the impeller and vaned diffuser (Figure 7(b)). 

Similar hub-side injection in the vaneless region was found to 

be less effective. In addition to steady shroud-side air 
NASA/TM—2012-217280 8

injection, Skoch found that tubes protruding into the vaneless 

region through the shroud surface improved stability margin 

by 4.4 percent. Both the injectors and stall-control-tubes have 

an associated pressure ratio penalty of approximately

1 percent in the centrifugal compressor, depending upon the 

injection flow rate or control-tube immersion.

T700 Engine Experiments

The stall control technologies described above were 

subsequently applied in a T700 engine tests. The T700 

turboshaft engine was selected for the engine demonstration of 

stall control due to its important role in powering the Army 

helicopter fleet. The compressor of the T700 engine has five 

and a half axial stages, the first three of which have variable 

inlet-guide-vanes/stators, followed by a single-stage

centrifugal. Experiments were carried out in the NASA Engine 

Components Research Laboratory (ECRL, Figure 8), and 

involved stall control in both the axial and centrifugal stages. 

The axial and centrifugal compressors could be throttled 

independently by preferential back-pressuring using externally 

controllable inflow of high-pressure air.

Coanda injectors of the type demonstrated by Strazisar et al.

(Ref. 22) were installed ahead of Rotors 1, 2 and 5 of the axial 

compressor. The injection air was supplied from either an 

external high pressure source with temperature control or by 

recirculation from within the engine. Injectors at the Rotor 1 

and 2 locations were supplied by air extracted downstream of 

the last axial stage, while Rotor 5 injectors were supplied from 

air extracted downstream of the centrifugal compressor stage.

In addition to the air injectors in the axial compressor, 

retractable, solid, stall control tubes were installed at multiple 

locations around the circumference in the vaneless space 

ahead of the centrifugal compressor diffuser.

Example results for tip injection in the axial compressor are 

shown in Figure 9 where range extension is indicated for 

operation at 80 percent speed. In the upper chart, three of the 

six available injectors ahead of Rotor 1 were active as the 

axial compressor was throttled. Tip injection provided a

3.5 percent increase in flow range, while maintaining stall 

pressure ratio.

The ECRL test facility provides robust data collection 

capability. In addition to versatile injector and engine 

controls, the ARL/NASA T700 engine is heavily instrumented. Total pressures and temperatures are measured 

upstream and downstream of both the axial and centrifugal 

compressors. The leading and trailing edge of each axial 

blade row, rotors and stators are instrumented with steadystate casing static-pressures. High-response pressure 

measurements are collected upstream of each axial rotor as 

well. In the near future, a light probe system will be added to 

measure blade vibration during stall and surge. The diffuser 

is similarly heavily instrumented with steady-state and highresponse pressure instrumentation.

URANS Simulations of Stall Inception and Control

In addition to the T700 engine testing, the ASCED program 

included a compressor simulation element. Unsteady Reynoldsaveraged Navier-Stokes (URANS) equation simulations have 

been used to elucidate the causal mechanisms and associated 

flow physics leading to stall—that is, stall inception—and the 

rotor-passage-level impact of the discrete casing endwall 

injection used for stall control. The simulations complement 

component level (see Weigl et al. (Ref. 25)) and engine level 

(above) experiments aimed at understanding stall inception and 

were intended to provide guidance in the development and 

testing of tip injection stall control technology. In this section, a 

brief review of the simulation activities is provided, including 

their application to support T700 engine experiments.

Figure 8.—Photograph of the NASA Engine Component 

Research Laboratory with T700 engine and stall control 

injection air lines installed.

Figure 9.—Demonstrated stall control in five-stage axial 

compressor in T700 engine at (80 percent) operating speed, 

showing flow range extension achieved by injection ahead of 

Rotor 1. Normalized axial compressor pressure ratio

Normalized inlet corrected flow (lbm/s)

0.65 0.70 0.75 0.80 0.85 0.90 0.95 1.00 1.05

0.80

0.90

1.00

1.10

0.65 0.70 0.75 0.80 0.85 0.90 0.95 1.00 1.05

0.80

0.90

1.00

1.10

Three Rotor-1 injectors active

No injectors active (baseline)

3.5%
NASA/TM—2012-217280 9

Initially steady axisymmetric simulations (Hathaway and 

Strazisar (Ref. 26)) with and without casing tip injection 

modeled, indicated that stall was initiated when the rotor tip 

section exceeded a critical incidence (or max. diffusion factor), 

and that injection velocity and number of injectors (specifically 

circumferential coverage) were key parameters for effective 

range extension (and cf. Suder et al. (Ref. 21)). Such axisymmetric simulations did not admit the non-engine order, nonaxisymmetric spatio-temporal flow field variations associated with 

stall inception and stall; indeed, modeling stall and its inception 

requires time- and CPU-intensive 3D, unsteady full-annulus 

simulations. Full-annulus simulations were thus conducted for 

both axial-flow and centrifugal compressors with a principal 

objective to further understand the causal fluid mechanisms of 

rotating stall (Hathaway et al. (Ref. 27) and Chen et al. 

(Ref. 28)). Subsequent simulations (Chen et al. (Ref. 29)) of the 

axial compressor as it was throttled to stall demonstrated the 

ability to simulate the initiation of flow instabilities and their 

subsequent growth into a fully developed rotating stall, without 

introduction of an embryonic disturbance.

The beneficial impact of steady tip injection has been 

demonstrated in recent simulations in which modeling of the 

discrete tip injection stall control technology has been 

included (Chen et al. (Ref. 30)). The analyses have indicated 

that, without stall control, the initial regions of low or reversed 

axial velocity (disturbance cells) first develop in the mid-span 

region of the rotor suction surfaces (Figure 10(a)), most likely 

due to the increased incidence and diffusion factor values for 

this region as it is throttled beyond the stable operating point. 

These disturbance cells migrate toward the tip region while 

moving downstream. They increase in size as they move to the 

more highly loaded tip region. When tip injection is of 

sufficient magnitude to stabilize the compressor, the 

disturbance cells are successively eliminated by the injectors 

(Figure 10(b)). Some disturbance cells were eliminated after 

passing through only one injector, while others must pass

through multiple injectors before being eliminated.

While stall control by tip injection has been demonstrated to 

extend operation beyond the baseline (noninjection) stability 

boundary, the achieved stall range extension is limited by the 

magnitude and circumferential extent of injection and the 

ability of the lower spans to accommodate the increase stall 

range capability. As described, the compressor with tip 

injection is throttled ever deeper beyond the noninjection stall 

point, the lower spans subsequently exceed their loading 

capability (high incidence and diffusion factor) resulting in 

local reverse flow pockets that centrifuge out to the rotor tip 

coalescing into stall cells, which eventually grow and spill 

flow ahead of the rotor leading edge, eventually leading to 

stall (see Figure 11).

Both steady axisymmetric and unsteady full-annulus 

simulations of the five and a half stage axial compressor of the 

T700 engine were conducted in support of the ASCED project 

(Hathaway et al. (Ref. 27)). Efforts to simulate the centrifugal 

compressor stage were also initiated with an eye toward 

simulating the complete T700 axicentrifugal compression 

system, both with and without stall control technology. These 

computationally intensive simulations involved thousands of 

processors and CPU-months worth of computations to 

determine the stall boundaries. The investment proved 

impractical for generic engine simulation in which bleed 

flows, secondary flows, and variable geometry schedules 

Figure 10.—Pictures of computed instantaneous size and location of local regions of reversed flow (disturbances) in a tip-injection 

stabilized rotor operating beyond the baseline stability boundary, showing (a) radial transport of disturbances from mid-span to tip 

region and (b) clean-up of disturbances in tip region by injected flow (Chen et al. (Ref. 30)).

b. Disturbance reduction via discrete tip injectors at 

0.8 rotor pitch time increments per frame in 

stabilized operation beyond baseline stability 

boundary.

a. Radial migration of the disturbance at 0.4 rotor 

pitch steps per frame in the rotor in stabilized 

operation beyond baseline stability boundary.
NASA/TM—2012-217280 10

strongly impact operability. While the URANS simulations 

continue to serve a role in elucidating the flow physics of stall 

and its inception, more rapid and less CPU intensive 

computations are required for rapid simulation of full-annulus, 

unsteady compressor flow fields.

Unsteady Actuator-Duct Solver for 

Nonaxisymmetric Flows

To this end, a more recent effort involves development of a 

3-D, unsteady, nonaxisymmetric actuator-duct model for rapid 

simulation of multistage transonic fans/compressors with 

nonaxisymmetric flow fields of low azimuthal wave number. 

The model is intended for the simulation of the steady and 

dynamic response of multistage and multi-spool compression 

systems subjected to generalized inlet circumferential 

distortions and other nonaxisymmetric effects, including 

discrete endwall injection.

The Navier-Stokes equations are low-pass filtered 

(Figure 12(a)) in the circumferential direction, resulting in 

governing equations which admit only long wavelength 

circumferential nonuniformities (e.g., wave numbers less than 

blade count) and which cannot resolve the short wavelength 

nonuniformities at the blade-passage level (cf. Xu et al.

(Ref. 31)). The filtered equations are similar in form to the 

Navier-Stokes equations, with the addition of body force terms 

for the blade forces and an additional transport equation for 

the kinetic energy associated with the shorter (filtered) 

wavelength distortions. The kinetic energy of the short 

wavelength distortions is convected by the mean flow, 

transported by accelerations due to flow work, produced by 

viscous dissipation and dilatation flow, produced by power

(dissipation) due to blade forces, and produced or destroyed 

by a term which scatters energy to/from the deviation and 

filtered flow fields. The Harmonic Balance technique (Thomas 

et al. (Ref. 32)) is used to resolve the circumferential 

derivative of the θ-flux term as in the earlier compressor 

model for generalized inlet distortions by Welch (Ref. 33),

resulting in an additional source term that couples 2N+1 

“throughflow” solutions at equally spaced θ-locations, where 

N is the maximum wave number (harmonic order) of the 

solution. 

Figure 11.—Computed streaklines of clearance flow during 

near-stall operation, showing spillage of clearance flow past 

rotor leading edge (Chen et al. (Ref. 30)).

Figure 12.—(a) Schematic showing low-pass filtering at local circumferential position; and (b) computed 

contours of static pressure for NASA Rotor 35 at 100 percent speed, steady-state operation (axisymmetric 

case).

direction of rotation

spillage

L.E.

T.E

.

rΩ

2

θ f θ

∆

+

2

θ f θ∆ −

∆θ f

Filter

window

at θ

θ upper

limit

lower

limit

rΩ

2

θ f θ

∆

+

2

θ f θ∆ −

∆θ f

Filter

window

at θ

θ upper

limit

lower

limit

a. Schematic of low-pass filter b. Computed static pressure contours 
NASA/TM—2012-217280 11

The blade forces must be supplied to the computation, and 

are currently set equal to the blade forces that would exist if 

the flow were axisymmetric at the local (θ) conditions of the 

nonaxisymmetric flow. Chima (Ref. 34), Hale et al. (Ref. 35),

and Longley (Ref. 36) use similar modeling approaches, with 

principal differences being the blade force specification. In 

Chima’s case, the blade forces are calculated in separate 

steady-state 3-D Navier-Stokes calculations, whereas in Hale’s 

case, forces are obtained from axisymmetric throughflow 

modeling using a streamline curvature method. The present 

effort is intended for modeling multistage machines, which 

like the T700 might have order 10 blade rows; therefore, 

rather than computing the blade forces directly, the body 

forces are obtained approximately (cf. Longley (Ref. 36)) at 

the blade row level by assuming (i) a specified turning 

schedule which turns the incoming flow to a local blade 

departure angle based on the blade geometry and the 

computed short wavelength distortion intensity (aeroblockage); and (ii) the blade force decomposition described in 

Welch and Larosiliere (Ref. 37), which relates the local aeroloading corresponding to the local flow turning to the bladeforce components.

The nonaxisymmetric actuator-duct modeling effort 

represents a work in progress. A preliminary result from an 

inviscid (Euler) calculation of NASA Rotor 35 is shown in 

Figure 12(b). The coarse grid used for rotor alone and rotor 

and stator computations is indicated as well. Upon completion, 

the model is intended to enable efficient simulation of the 

dynamics of compressors with nonaxisymmetric flow field. 

The planned next steps include assessment and validation of 

the model using known axisymmetric performance data (e.g., 

NASA stage 35 and CC3 centrifugal compressor) followed by 

simulation of steady-state and dynamic T700 engine 

performance in support of the stall control effort.

Conclusions

An overview of on-going compressor research at NASA 

GRC which is intended to address technical challenges for 

future rotorcraft engines was provided. With particular 

attention given to the variable rotor requirements of LCTR 

future applications, the technical challenges for the compressor are associated with low corrected flows in the aft stages, 

high compressor discharge temperatures, and compressor 

stability management. The on-going research described 

includes modeling and experimental efforts that address 

aerodynamic challenges associated with low corrected flow aft 

stages and compressor stability. Described modeling efforts 

are intended to enhance future predictive and assessment 

capability for multistage compressors, at both the cycle deck 

and component simulation levels. The T700 testing in the 

Engine Components Research Laboratory is demonstrating the 

applicability of stall control approaches in the engine, and 

related URANS simulations are elucidating the flow physics 

of stall and its inception. Finally, the 3-year NRA contract is 

underway to develop a state-of-the-art centrifugal compressor 

stage with advanced aerodynamics which is to be tested in the 

NASA Small Engine Component Test Facility. The intent is 

that this new state-of-the-art rotorcraft-relevant compressor 

stage will serve as an experimental test vehicle for attainment 

of industry/academia/government consensus data sets needed 

for code validation and to fill gaps in the knowledge base.

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SP–36, Chapter 10, 1965, pp. 297–310.
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14. Smith, L.H., Jr., “The Effect of Tip Clearance on the Peak 

Pressure rise of Axial Flow Fans and Compressors,” 

ASME Symposium on Stall, 1958, pp. 149–152.

15. Koch, C.C. and Smith, L.H., Jr., “Loss Sources and 

Magnitudes in Axial-Flow Compressors,” J. of Eng. for 

Power, Vol. 98, 1976, pp. 411–424.

16. 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.

17. Skoch, G.J., Prahst, P.S., Wernet, M.P., Wood, J.R., and 

Strazisar, A.J., “Laser Anemometer Measurements of the 

Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller,” 

NASA TM–107541, June 1997.

18. Wernet, M.P., Bright, M.M., and Skoch, G.J., “An Investigation of Surge in a High-Speed Centrifugal Compressor 

Using Digital PIV,” J. Turbomachinery, Vol. 123, Apr. 

2001, pp. 418–218. 

19. Larosiliere, L.M., Skoch, G.J., Prahst, P.S., “Aerodynamic Synthesis of a Centrifugal Impeller using CFD and 

Measurements,” AIAA–97–2878, July 1997; also NASA 

TM–107515 and ARL–TR–1461, July 1997.

20. Adamczyk, J.J., Hansen, J.L., and Prahst, P.S., “A Post 

Test Analysis of a High-Speed Two-Stage Axial Flow 

Compressor,” ASME GT2007–28057, May 2007.

21. Suder, K.L., Hathaway, M.D., Thorp, S., Strazisar, A.J., 

Bright, M.M., “Compressor Stability Enhancement Using 

Discrete Tip Injection,” J. Turbomachinery, Vol. 123 (1), 

Jan. 2001, pp. 14–23.

22. Strazisar, A.J., Bright, M.M., Thorp, S., Culley, D.E., 

Suder, K.L, “Compressor Stall Control through Endwall 

Recirculation,” ASME GT2004–54295, June 2004.

23. Skoch, G.J., “Experimental Investigation of Centrifugal 

Compressor Stabilization Techniques,” J. Turbomachinery, Vol. 25, Oct. 2003, pp. 704–713.

24. Skoch, G.J., “Experimental Investigation of Diffuser Hub 

Injection to Improve Centrifugal Compressor 

Stability,” J. Turbomachinery, Vol. 127, Jan. 2005, pp. 

107–117.

25. Weigl, H.J., Paduano, J.D., Frechette, L.G., Epstein, A.H., 

Greitzer, E.M., Bright, M.M., and Strazisar, A.J., “Active 

Stabilization of Rotating Stall and Surge in a Transonic 

Single Stage Axial Compressor,” J. Turbomachinery, Vol. 

120 (4), 1998, pp. 625–636.

26. Hathaway, M.D. and Strazisar, A.J., “Impact of Discrete 

Tip Injection on Stabilization of a Transonic Compressor 

Rotor,” Proc. 21st Army Science Conference, Norfolk, 

Virginia, June 15–17, 1998.

27. Hathaway, M.D., Chen, J., Webster, R., G.P. Herrick, 

“Unsteady Simulation of the Stall Inception Process in the 

Compression System of a US Army Helicopter Gas 

Turbine Engine – Final Year Progress,” DoD High Performance Computing Modernization Office sponsored 

“2005 Users Group Conference”, Nashville, TN, June 27–

30, 2005.

28. Chen, J.-P., Chen, J.P., Webster, R.S., Hathaway, M.D., 

Herrick, G.P., and Skoch, G.J., “Numerical Simulation of 

Stall and Stall Control in Axial and Radial Compressors,” 

AIAA 2006–418, Jan. 2006.

29. Chen, J.-P., Hathaway, M.D., Herrick, G.P., “Prestall 

Behavior of a Transonic Axial Compressor Stage via 

Time Accurate Numerical Simulation,” J. Turbomachinery, Vol. 130, Oct. 2008, p. 041014.

30. Chen, J.P., Johnson, B., Hathaway, M.D., Webster, R.S., 

“Flow Characteristics of Tip-Injection on Compressor

Rotating Spike via Time-Accurate Simulation,” (accepted 

for publication in the 2009 AIAA J.).

31. Xu, L., Hynes, T.P. and Denton, J.D., “Towards Long 

Length Scale Unsteady Modelling in Turbomachines”, 

Proc. IMechE, Pt A: J. Power and Energy, Vol. 217, 

2003.

32. Thomas, J.P., Dowell, E.H., and Hall, K.C., “A Harmonic 

Balance Approach for Modeling Three-Dimensional 

Nonlinear Unsteady Aerodynamics and Aeroelasticity,” 

IMECE–2002–32532, Nov. 2002.

33. Welch, G.E., “Determination of Critical Sector Angle of 

Inlet Swirl Distortion,” AIAA 2007–5057, July 2007.

34. Chima, R.V., “A Three-Dimensional Unsteady CFD 

Model of Compressor Stability,” ASME GT2006–90040, 

May 2006.

35. Hale, A., and O’Brien, W.O., 1998, “A ThreeDimensional Turbine Engine Analysis Compressor Code 

(TEACC) For Steady State Inlet Distortion,” J. Turbomachinery, Vol. 120, pp. 422–430.

36. Longley, J.P., “Calculating Stall and Surge Transients,” 

ASME GT2007–27378, May 2007.

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Description of Wave Rotor Flow,” AIAA–97–3144, July

1997.

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Technical challenges of compressors for future rotorcraft engines are driven by engine-level and component-level requirements. Cycle 

analyses are used to highlight the engine-level challenges for 3000, 7500, and 12000 SHP-class engines, which include retention of 

performance and stability margin at low corrected flows, and matching compressor type, axial-flow or centrifugal, to the low corrected 

flows and high temperatures in the aft stages. At the component level: power-to-weight and efficiency requirements impel designs with 

lower inherent aerodynamic stability margin; and, optimum engine overall pressure ratios lead to small blade heights and the associated 

challenges of scale, particularly increased clearance-to-span ratios. The technical challenges associated with the aerodynamics of low 

corrected flows and stability management impel the compressor aero research and development efforts reviewed herein. These activities 

include development of simple models for clearance sensitivities to improve cycle calculations, full-annulus, unsteady Navier-Stokes 

simulations used to elucidate stall, its inception, and the physics of stall control by discrete tip-injection, development of an actuator-ductbased model for rapid simulation of nonaxisymmetric flow fields (e.g., due inlet circumferential distortion), advanced centrifugal 

compressor stage development and experimentation, and application of stall control in a T700 engine.

15. SUBJECT TERMS

Gas turbine engine; Turboshaft engine; Engine performance; Compressor aerodynamics; CFD 

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