Gerard E. Welch, Ashlie B. McVetta, Mark A. Stevens, and Samuel A. Howard

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

ASRC Aerospace Corporation, Cleveland, Ohio

Ali A. Ameri

The Ohio State University, Columbus, Ohio

Waiming To

University of Toledo, Toledo, Ohio 

Gary J. Skoch and Douglas R. Thurman

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

Variable-Speed Power-Turbine Research 

at Glenn Research Center

NASA/TM—2012-217605

July 2012
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientific and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Officer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of significant 

scientific and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Gerard E. Welch, Ashlie B. McVetta, Mark A. Stevens, and Samuel A. Howard

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

ASRC Aerospace Corporation, Cleveland, Ohio

Ali A. Ameri

The Ohio State University, Columbus, Ohio

Waiming To

University of Toledo, Toledo, Ohio 

Gary J. Skoch and Douglas R. Thurman

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

Variable-Speed Power-Turbine Research 

at Glenn Research Center

NASA/TM—2012-217605

July 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

68th Annual Forum and Technology Display (Forum 67)

sponsored by the American Helicopter Society (AHS)

Fort Worth, Texas, May 1–3, 2012
Acknowledgments

This effort is conducted under the NASA Fundamental Aeronautics Program, Subsonic Rotary Wing Project. The authors thank 

M. Suchezky of Williams International, L.L.C. and A. Ford, M. Bloxham, S. Gegg, and E. Turner of Rolls-Royce North American 

Technologies and Rolls-Royce Corporation for their contributions, and insights shared, during the VSPT contracts referenced 

herein. Special thanks to R.J. Boyle (retired from NASA) and Dr. M.D. Hathaway (NASA) for their early contributions in setting 

the VSPT research vector. Thanks to C. Snyder (NASA) for providing LCTR concept-engine performance numbers.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

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 

Space Administration.

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. 

This report contains preliminary findings, 

subject to revision as analysis proceeds.
NASA/TM—2012-217605 1

Variable-Speed Power-Turbine Research at Glenn Research Center

Gerard E. Welch, Ashlie B. McVetta, Mark A. Stevens, and Samuel A. Howard

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Paul W. Giel

ASRC Aerospace Corporation

Cleveland, Ohio 44135

Ali A. Ameri

The Ohio State University

Columbus, Ohio 43210

Waiming To

University of Toledo

Toledo, Ohio 43606

Gary J. Skoch and Douglas R. Thurman

U.S. Army Research Laboratory

Glenn Research Center

Cleveland, Ohio 44135

Abstract

The main rotors of the NASA Large Civil Tilt-Rotor 

(LCTR) notional vehicle operate over a wide speed-range, 

from 100 percent at takeoff to 54 percent at cruise. The 

variable-speed power turbine (VSPT) offers one approach by 

which to effect this speed variation. VSPT aerodynamics 

challenges include high work factors at cruise, wide (40° to 

60°) incidence-angle variations in blade and vane rows over 

the speed range, and operation at low Reynolds numbers. 

Rotordynamics challenges include potential responsiveness to 

shaft modes within the 50 percent VSPT speed-range. A 

research effort underway at NASA Glenn Research Center, 

intended to address these key aerodynamic and rotordynamic 

challenges, is described. Conceptual design and 3-D multistage RANS and URANS analyses, conducted internally and 

under contract, provide expected VSPT sizing, stage-count,

performance and operability information, and maps for system 

studies. Initial steps toward experimental testing of incidencetolerant blading in a transonic linear cascade are described, 

and progress toward development/improvement of a simulation capability for multistage turbines with low Reynolds 

number transitional flow is summarized. Preliminary rotordynamics analyses indicate that viable concept engines with 

50 percent VSPT shaft-speed range. Assessments of potential 

paths toward VSPT component-level testing are summarized.

Nomenclature

AN2 product of annulus area and shaft-speed squared

Cp

p0,1−p

1

2ρU2 , local pressure coefficient

Cptot

p0,1−p0

p0,1−p2

, local total-pressure coefficient

cx axial chord

h0 total specific enthalpy

i, ides incidence, design incidence

k turbulent kinetic energy

kl laminar kinetic energy

N power-turbine shaft speed

N* N/N100%, fraction of 100 percent speed

Nu Nusselt number

PRTT overall total-pressure ratio

p0, p total and static pressure

Recx Reynolds number based on axial chord

s blade pitch, or specific entropy

T0 total temperature

TRTT overall total-temperature ratio

Tu turbulence intensity

u (ux, uθ, ur), absolute velocity

U blade speed at pitchline

Xpk-pk vibration amplitude, peak-to-peak

Z s

cx

ρux(uθ,1−uθ,2)

p0,r,1−p2

, Zweifel loading parameter

α, β absolute and relative flow angles

ηTT adiabatic efficiency (total-to-total)

ρ density

ψ ∆h0/U2

, work factor

φ ux/U, flow coefficient

ω p̅

0,1−p̅0,2

p̅

0,1−p̅2

, loss coefficient (area-averaged)

ω specific dissipation rate of turbulence
NASA/TM—2012-217605 2

Subscripts

c corrected to standard day

r relative conditions

1, 2 Blade-row inlet, blade-row exit

4.5 power turbine inlet

Introduction

The vertical takeoff and landing (VTOL) and Mach 0.5

cruise capability of the NASA Large Civil Tilt-Rotor (LCTR,

(Refs. 1 and 2) Figure 1) is envisaged to increase throughput in 

the national airspace (Refs. 1 and 3). A key challenge of the 

LCTR is the variation in main-rotor tip-speed from 650 ft/s

(100 percent N*

) at sea-level VTOL to 350 ft/s (54 percent N*) 

at 28 k-ft cruise as required to minimize mission fuel-burn. 

The variable-speed power-turbine (VSPT), driving the rotor 

through a fixed gear-ratio transmission, provides one approach 

for effecting this wide speed variation. The VSPT approach is 

used in the V-22 where the output speed of the AE1107 engine 

varies speed in the range 80 percent < N* ≤ 100 percent

(Ref. 3). The LCTR mission evidently widens the operational 

speed range of the main rotor, drive system, and VSPT powershaft (54 percent ≤ N*

≤ 100 percent) considerably.

Key Technical Challenges

The VSPT research effort at NASA Glenn Research Center

(GRC) is intended to address key aerodynamic and mechanical (rotordynamic) challenges of the VSPT. The aerodynamic 

challenges include attainment of high turbine efficiency at 

high work factor, management of loss levels over a large (60°) 

incidence variation in all blade rows associated with the shaftspeed change, and operation at low Reynolds numbers with 

attendant sensitivity of transitional flow (Ref. 4). The loss 

levels in the transitional flow fields of low-pressure turbines 

(LPTs) operating at cruise altitudes (with Re < 100 k) are 

known to be impacted strongly by wake-induced unsteadiness

(Ref. 5) as well. The mechanical challenges are associated 

with the required avoidance or management of responsive 

shaft, blade, and casing modes at critical speeds within the 

operational speed range of the VSPT shaft (Ref. 6).

While the VSPT approach to speed change is the subject of 

this paper, the main-rotor speed variation could alternatively 

be accomplished by using a two-speed (reduction gear-ratio) 

transmission (Ref. 7) (shifting) with a fixed-speed power 

turbine. The turbine of the two-speed transmission option 

shares the aerodynamic challenges associated with low-Re 

number operation and the need to avoid residence at shaft 

critical speeds during the shift maneuvers (Ref. 8). The 

distinguishing turbine benefits of the two-speed shifting option 

are associated with the ability to run the power turbine, 

throughout the mission, at the higher shaft speed of takeoff 

(100 percent N*

) with minimal incidence variation. Operation 

at the takeoff shaft-speed with design-intent air angles enables 

Figure 1.—NASA Large Civil Tilt-Rotor notional vehicle (Ref. 1).

higher cruise and takeoff efficiency levels, achievable at lower 

design-point work factors, in lower weight, reduced stagecount turbine designs.

Research Approach

The VSPT technical challenges have motivated a research

(Ref. 4) effort at GRC related to VSPT aerodynamics, rotordynamics, and component-level testing. These research 

activities were conducted internal to GRC and through 

contracted efforts with industry.

The aerodynamics research elements include:

i. VSPT requirements and conceptual design. Execute conceptual design of a VSPT to meet LCTR mission/engine

requirements at the meanline level, to establish representative 

blade row performance levels and VSPT performance maps.

ii. Detailed design of incidence-tolerant blading. Conduct 3-

D aero-design (computational) of incidence-tolerant rotor, 

vane, and EGV blade shapes using flow angles from the 

meanline analyses. 

iii. Transonic cascade testing of 2-D sections. Test 2-D 

sections of the 3-D VSPT incidence-tolerant blading in GRC 

transonic linear cascade at LCTR-relevant Mach numbers, 

Reynolds numbers, incidence angle variations (>55°).

iv. Transition sub-model assessment/implementation. Develop a validated 3-D unsteady RANS simulation capability 

for unsteady low-Re transitional flows in multistage VSPT 

blade rows.

The rotordynamics research element addresses mechanical 

challenges associated with operation of a VSPT, with intent to 

identify potential constraints imposed on the aerodynamic 

design by mechanical issues associated with variable speed—

for example, growth in shaft radius to meet torque capacity 

requirements and to manage responsiveness to shaft critical 

speeds (rotordynamics).
NASA/TM—2012-217605 3

Finally, the component test element is to assess potential inhouse and external test vehicles for component level VSPT 

experimentation.

Content

An overview of progress to-date on the research elements 

outlined above is provided herein. The paper includes sections 

related to VSPT conceptual aerodynamic design and analysis, 

experimental testing of incidence-tolerant blading, computational methods, rotordynamics, and an assessment of potential 

VSPT component test paths. A summary is provided at the 

conclusion of the paper.

VSPT Conceptual Aerodynamic

Design and Analysis

The conceptual design and analysis of VSPTs for the LCTR 

mission was conducted by NASA and through NASA Research and Technologies for Aerospace Propulsion Systems

(RTAPS) contracts with Williams Int., L.L.C. (Ref. 9) and 

Rolls-Royce North American Technologies (Refs. 10 and 11).

The intent of the RTAPS efforts were to develop a conceptual

design of the VSPT for the LCTR application to a level 

required to emerge the key technical challenges—both 

aerodynamic and rotordynamic—to identify research required 

to overcome these challenges, and to define a path forward, 

and the test-plan, for a VSPT component-level test. 

NASA, Williams International (WI) (Ref. 9) and RollsRoyce North American Technologies (RR-NAT) (Ref. 10)

arrived at a 4-stage VSPT design to meet the LCTR mission 

and engine cycle provided by the NASA engine performance 

group. Rolls-Royce carried out a second effort using a higher 

mass-flow-rate, growth AE1107C engine cycle, which led to a 

3.5-stage configuration (Ref. 11). In all studies, the VSPT 

requirements were met without the use of variable geometry.

The turbine requirements of the VSPT concept design effort 

are provided below, followed by key findings from each 

conceptual design and aerodynamic analysis efforts by NASA 

and industry.

VSPT Requirements

The NASA LCTR notional vehicle (Figure 1) is envisaged 

to be a 90 passenger aircraft with 100,000 lbf takeoff gross 

weight. There are two engines per nacelle (and main rotor), 

currently conceived to be of the 7500 shp-class level. Two key 

LCTR mission points are 2 k-ft takeoff/hover and 28 k-ft, 

Mach 0.5 cruise. At takeoff, the main rotor and VSPT operate 

at 100 percent N*

, while at cruise the rotors and VSPT are 

slowed to 54 percent N*

. The LTRC engine requirements were 

established with the NASA engine performance group

(Ref. 12). The corresponding VSPT requirements for the key 

takeoff and cruise operating points are provided in Table I.

TABLE I.—VSPT REQUIREMENTS AT KEY FLIGHT 

POINTS OF LCTR MISSION (REF. 12)

Flight point

Altitude

VSPT shaft-speed (N/N100%)

Main-rotor tip-speed

Takeoff 

2k-ft

100%

650 ft/s

Cruise

28 k-ft

54%

350 ft/s

Power, shp 4593 2328

Mass flow rate, lbm/s 22.03 12.22

Specific power (Btu/lbm) 147 135

Inlet temperature (T0,4.5), R 2204 1812

Inlet pressure (p0,4.5), psia 58.0 26.76

PRTT 4.04 5.34

Corrected flow, lbm/s 11.51 12.54

Corrected speed (Nc/Nc,100%), % 102.3 60.8

Aft-stage unit-Re (in-1)* 70,000/in. 45,000/in.

*Based on static conditions at last stage rotor with exit relative Mach 

number of 0.7.

NASA Aerodynamic Design and Aero Analysis

VSPT Conceptual Design

The NASA conceptual aero-design approach for a 4-stage 

VPST of the LCTR application was outlined earlier (Refs. 4

and 13). The design air-angles were set at the cruise operating 

condition (54 percent N*

) where Reynolds numbers were

lowest and work factors (∆h0/U2

) were highest. Maximum 

vane and blade-row incidence levels were as high as negative 

60° at takeoff (100 percent N*

). The turbine flow path, number 

of stages, and design air angles were determined using F. 

Huber’s meanline design and off-design codes, which are 

constituents of the AFRL Turbine Design and Analysis 

System (TDAAS) (Ref. 14). The aerodynamic loading levels 

of the vanes (Zweifel coefficient) were set near unity so as to 

be consistent with operation with transitional flow (Ref. 15)

and required incidence tolerance. The stage reaction levels 

were set near 0.45 and flow coefficients, ux/U, at cruise were 

set near unity. The stage work splits were based on trades 

between optimum efficiency and management of maximum 

turning per stage. The rotors were tip-shrouded and leakage 

flows were neglected. Huber’s off-design code (Ref. 14)

(meanline) was used to assess off-design operation.

The design- (54 percent N*

, cruise) and off-design (100

percent N*

, takeoff) performance levels are spotted on the 

modified Smith Chart of Figure 2. AN2 levels, 47.7×109

rpm2

∙in2 at 15,000 rpm takeoff speed, were based on strengthof-materials for anticipated turbine materials at the operating 

conditions of the LCTR VSPT. The average work factors from 

industry conceptual designs (Refs. 9, 10, and 16) are shown as 

well. The impact of adding a stage, from 3-stages to 4-stages, 

on efficiency corresponding to design work factor is evident. 

The gained VSPT efficiency at the higher stage count would 

be expected to come at the cost of increased weight to the 

VSPT. The sensitivity of the LCTR to the turbine weight and 

efficiency makes the stage count selection a deep trade that 

will aim to optimize the VSPT efficiency and weight, for 

optimal life, cost, and fuel burn.
NASA/TM—2012-217605 4

Figure 2.—Modified Smith chart showing meanline predicted design-point 

efficiency as a function of work factor for 3- and 4-stage VSPT meanline 

designs (Refs. 9, 10, 13, and 16) compared to Smith (Ref. 17) and Oates 

(Ref. 18) correlations and AMDCKO methodology (Refs. 19 to 21). Calculated 

performance levels at the off-design takeoff (100 percent N*

) are also shown.

The off-design efficiency level, corresponding to takeoff 

operation with strongly negative incidence (e.g., 

–60°) in all blade rows, is predicted (Figure 2) by meanline

analyses (Refs. 4, 9, 10, 13, and 16) to be higher than that at 

design-point where work factor is highest (Refs. 4 and 13). 

The overall loss levels are lower at the reduced work factors 

and aerodynamic loading levels, Z, of takeoff. The profile loss 

component—associated with boundary-layer and shocks—at 

negative incidence is near that of the higher loaded design air 

angles (see Ref. 4) whereas the secondary-flows and leakage 

flows (3-D) predicted at the lower aerodynamic loading levels 

of off-design are substantially lower than at the on-design 

cruise condition. The impact of this counterintuitive efficiency 

increase as the turbine unloads biases the selected VSPT 

design speed toward the cruise speed (54 percent N*

).

3-D Blading Design

The predicted loss levels at off-design are strongly dependent 

on the loss correlations in the meanline tool used (Refs. 9, 10, 

and 13). Therefore a 3-D CFD study of a representative embedded 1.5-stage (Rotor 1, Stator 2, and Rotor 2, Figure 3) of the 4-

stage NASA VSPT was conducted to verify the efficiency 

versus speed trend of the meanline code held when 3-D flow 

structures (see Figure 3 and Figure 4) were present. The 3-D 

aero-design of the blading for the 1.5-stage (R1/S2/R2) of the 

present study was described in detail earlier (Ref. 13). R1 and 

R2 were shrouded, and leakage flows were neglected. With

Figure 3.—Computed contours of axial velocity and streamlines in R1 and R2 at the 100% N* off-design take-off point 

(reproduced from Ref. 22). 

design air angles from the meanline design code, the AFRL 

Turbine Design and Analysis System (TDAAS (Ref. 14)) was 

used to design, optimize, and stack 2-D airfoil sections, at hub, 

midspan, and tip, into a 3-D blade geometry. The chordwiselocation of maximum loading was influenced by the findings 

of Gier et al. (Ref. 15) concerning efficiency, Re-lapse, and 

sensitivity to incidence in low pressure turbines designed for

0.76

0.78

0.8

0.82

0.84

0.86

0.88

0.9

0.92

0.94

0.96

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5

Overall adiabatic efficiency

Average stage work factor, ∆h0/U2

Smith (1965)

LPT Oates

AMDCKO methodology

NASA 3-stage

NASA 4-stage

D'Angelo 3-stages

D'Angelo 4-stage/tandem

Williams Int. L.L.C 4-stage.

Rolls-Royce NAT 3.5-stage w/ EGV

AMDCKO

1976 LPTs

(Oates)

Smith 

correlation

(1965)

40 to 60-deg.

negative incidence

At takeoff

100% N*

Improved Aero

R1

R2

S2
NASA/TM—2012-217605 5

(a) LEO RANS/mixing-plane

(b) LEO URANS TA/AP

Figure 4.—Computed contours of entropy at the blade-row 

exit planes from (a) LEO RANS/mixing-plane and (b) timeaveraged, average-passage (TA/AP) URANS calculations at 

design point (54 percent N*

, 28 k-ft cruise).

low-Re operation. The 2-D sections (provided in Ref. 13) were 

stacked on their center of gravity along a radial stacking axis. 

No dihedral was considered, though the benefits of 3-D 

aerodynamic design, in managing the secondary flow fields, is 

well documented (Ref. 5).

Efficiency/Speed Trend Verification 

Using the embedded 1.5-stage geometry, 3-D 

RANS/mixing-plane computations were conducted using the 

NASA SWIFT (Ref. 22) code at the cruise (54 percent N*

, 28 

k-ft) and takeoff (100 percent N*

, 2 k-ft) conditions. The 

meanline and 3-D simulations were found to agree in terms of 

the variation of efficiency with VSPT operating speed.

Consistent with meanline analyses, the blade-row loss levels 

were generally found to be lower at the off-design conditions 

where, though operating with 40° to 60° of negative incidence, 

the blade rows are unloaded. An exception to this general 

trend of higher efficiency at off-design was found in Rotor 2, 

attributable to the impact of the 3-D pressure-side cove vortex. 

This finding may push future rotor airfoil designs toward 

thicker sections that admit less of a cove separation. Enhancement of the meanline secondary loss model for rotors at 

extreme negative incidence, to account for increased loss due 

to such 3-D structures, may be warranted as well.

The agreement between 3-D RANS computations and the 

meanline both corroborated the incidence correlation of the 

meanline code—with the exception of R2 deficiency noted 

above—and supported the conceptual design approach that 

established the design speed at the lowest operating shaft speed 

(54 percent N*

, cruise). In practice, the specific VSPT design 

speed is expected to be strongly biased toward the cruise shaftspeed, and selected so as to minimize mission fuel burn.

Impact of Unsteadiness—URANS Simulations

The RANS/mixing-plane analysis (Ref. 13) was extended to 

3-D unsteady RANS (URANS) in an effort to assess the 

influence of unsteadiness associated with blade-row interaction on the design- and off-design aerodynamic performance 

of a VSPT for the NASA LCTR (Ref. 23). The extension was 

motivated by the potential impact on VSPT performance of 

flow physics not accounted for in mixing-plane analyses: 

additional entropy production associated with the unsteady 

forcing of the vortical structures (e.g., the pressure-side cove

R2 of Figure 3); spatiotemporally resolved loss production 

associated with wake transport and stretching (Ref. 24);

transport of streamwise vorticity to/from blade rows and 

associated impact of vorticity associated with secondary flow 

field of one blade row on the instantaneous and time-mean 

flow fields of downstream blade rows (Ref. 25); and, the 

deterministic vortical forcing of transitional boundary-layers 

by wakes from upstream blade rows (Refs. 26 and 27). The 

computations of the RANS/mixing-plane study (Ref. 13) were 

repeated using the commercially available WAND and LEO 

grid and flow solvers (Ref. 28) in both RANS/mixing-plane 

and sector-periodic URANS mode. Results from the 

RANS/mixing-plane and time-averaged, average-passage 

(TA/AP) URANS results are provided in Figure 4 where the 

blade row exit contours of the solutions are compared at the 

cruise conditions. 

 

R1

exit

R2

exit

S2

exit

R1

exit

R2

exit

S2

exit

 
NASA/TM—2012-217605 6

The agreement between the time-averaged/URANS and the 

RANS/mixing-plane solutions was excellent. The unsteadiness 

due to blade-row interaction was found to degrade work and 

efficiency levels in Rotor 2, which was most impacted by the 

upstream blade rows of the embedded R1/S2/R2 1.5-stage 

simulation. Quantification of the efficiency degradation 

(approximately 1 point) associated with unsteadiness, vis-à-vis 

comparison of efficiency predicted using the RANS/mixingplane and sector-periodic URANS computations, was unfortunately rendered inconclusive by a mixing-plane issue encountered (see Ref. 23). The trend of efficiency with changing 

VSPT shaft speed from the URANS simulation, from 54

percent N* (28 k-ft, cruise) to 100 percent N* (2 k-ft, takeoff), 

agreed with the corresponding trends from the two 

RANS/mixing-plane codes (SWIFT and LEO) and the 

meanline codes (Ref. 14) used for turbine design.

Williams International Aerodynamic Design and 

Analysis (Ref. 9)

VSPT Conceptual Design

In the Williams International (WI) RTAPS effort (Ref. 9), the 

shaft-speed at which to design the VSPT was determined 

through a sensitivity study of mission fuel burn versus the 

design speed. Turbine efficiencies were calculated using a

meanline tool proprietary to WI with loss correlations rooted in 

the open literature (consistent with AMDCKO (Refs. 19 to 21).

The results are reproduced in Figure 5 where it was found that 

the minimum fuel burn was achieved when the design point was 

set at approximately 70 percent N* (as opposed to the 54 percent

N* argued in the NASA work above). The trade between high 

and low flow coefficient ( φ = ux /U ) turbine designs—i.e., 

low/high turning versus high/low Mach numbers (Figure 5)—

high and low work factor, ∆h0/U2

, designs, and number of 

stages was also investigated. The conclusions were as follows

(Ref. 9): i.) there was no compelling evidence found at the 

meanline level to support a preferred selection of flow coefficient (Figure 5); ii.) of 3- and 4-stage designs, all 15000 rpm at 

100 percent N*

, a high work-factor (lower radius) 4-stage design 

was optimal in terms of mission fuel burn.

3-D Blading Design

A 3-D aero design was executed for the eight blade rows of 

the 4-stage VSPT, based on a meanline design at a 75 percent

N* (Ref. 9). Three 2-D sections were stacked for each of the 

eight blade rows. The 3-D sections were created using a WI

interactive design tool for blade geometry, structural analysis, 

2-D aero, and 3-D stacking. Design iterations were analyzed in 

a 3-D multistage Euler/RANS/mixing-plane solver 

(VORTEX) in Euler mode, and final designs were assessed 

using 3-D RANS/mixing-plane with k−ω turbulence modeling 

in VORTEX.

Figure 5.—Mission fuel burn as a function of VSPT design speed, showing 

insets for turbine designs at low and high flow coefficient (Ref. 9).
NASA/TM—2012-217605 7

The embedded third stage vane (S3) and blade (R3) blade 

were analyzed using the steady VORTEX flow solver to 

assess the impact of turbulence sub-model selection and blade 

thickness distributions. The shaft speed, N*

, was varied at a 

fixed stage pressure ratio. Regarding blade thickness, the 

intuition is that larger radius leading edges (and resultant blade 

thicknesses) afford incidence-tolerance, which must the traded 

against higher loss levels associated with higher through-flow 

Mach numbers. The study examined whether a thin blade 

(lower Mach numbers) might improve turbine efficiency over 

a traditional WI blade design, or whether admittance of a 

pressure-side cove separations at off-design, and subsequent 

centrifuging in thinned rotor blades (cf. NASA Rotor 2 of 

Figure 3), would outweigh gains. The thinner blades were not 

found to offer overall benefit over traditional blade thickness 

(area) schedules (Ref. 9).

Assessment of Performance/Speed Trend

A key finding of the WI study is reproduced in Figure 6, 

where efficiency of a 1.5-stage (R2/S2/R3) embedded stage, 

computed using the 3-D RANS/mixing plane VORTEX solver

is provided. The inlet conditions and pressure ratio of the 

embedded 1.5-stage of WI were held constant as speed was 

changed. The computed efficiency versus speed trend of the 

WI meanline tool and the 3-D computations are in qualitative 

agreement except that the roll-off in efficiency at negative 

incidence is more severe in the meanline than in the CFD. The 

efficiency at cruise is again found (see Fig. 2) to be lower than 

that at takeoff (extreme off-design). This finding corroborates

the embedded 1.5-stage results reported by NASA (Ref. 13

and 23), which are also plotted on Figure 6 The CFD results 

support the argument that the WI design-point speed (here 

75 percent of takeoff speed) might optimally be further 

reduced, giving stronger bias to cruise operation.

Rolls-Royce North American Technologies

Aerodynamic Design and Analysis

RR-NAT was awarded two NASA RTAPS contracts for an 

incidence-tolerant blade design effort (Ref. 10) and a VSPT 

study effort (Ref. 11) which paralleled the WI effort described 

above.

VSPT Conceptual Design

In the blade design contract (Ref. 10), meanline analyses

were used to determine that a 4-stage turbine was preferred to 

meet the LCTR engine requirements for the NASA mission 

provided. The meanline analysis was influenced by lessons 

learned from RR-NAT work done under subcontract to Boeing 

in a separate NASA contract (Ref. 29) and an early NASA 

concept design (Ref. 4). The VSPT was tailored to maintain 

matched stage flow coefficients near unity, effectively keeping 

axial Mach numbers below 0.5 through the four stages and 

balancing the stage work factors.

Figure 6.—Variation of computed efficiency with speed for Williams Int. embedded 

1.5-stage and NASA embedded 1.5-stage of 4-stage VSPTs for the LCTR 

mission.

78

80

82

84

86

88

90

92

94

96

30 40 50 60 70 80 90 100 110 120

Adiabatic efficiency, %

N*, percent

Williams Int. embedded 1.5-stage -meanline

Williams Int. embedded 1.5-stage -VORTEX RANS/mixing-plane

NASA embedded 1.5-stage -LEO URANS, time-averaged

NASA embedded 1.5-stage - SWIFT RANS/mixing-plane

NASA embedded 1.5-stage -Huber meanline - AFRL TDAAS
NASA/TM—2012-217605 8

Figure 7.—Computed midspan loss as a function of inlet 

relative air angle for Rotor 2 of 4-stage VSPT for LCTR 

application (Ref. 11), showing VSPT incidence-tolerant blade 

as inset.

Incidence-Tolerant Blade Design (Ref. 10) 

The second rotor (R2) of the RR-NAT 4-stage was selected 

as a representative embedded bladerow, both in terms of 

design turning and incidence range requirements between 

cruise and takeoff, for a 3-D blade design. Through-flow 

analysis was conducted using an RR-NAT axisymmetric 

streamline-curvature/radial-equilibrium solver. The 2-D 

streamline solution provided boundary conditions for subsequent 3-D design of R2. 

An incidence-tolerant midspan section was shape-optimized 

using the AIRFOILOPT code, which has been in use at RRNAT (formerly Allison Engine Company) for decades. A 

generic blade was first optimized at the cruise (design angles). 

New optimized airfoils shapes were then obtained at discrete 

inlet flow angle increments between cruise and takeoff 

(hover), with axial Mach number held constant. After traverse 

the full range of incidence, the midspan section was reoptimized at the cruise (design) condition. The loss bucket for 

the midspan section is provided in Figure 7 with the takeoff 

(hover, β1 = –2°) and cruise (β1 = +40°) inlet angles noted. A 

100 percent N* operating condition (β1 = –11°) corresponding 

to cruise altitudes is also noted. The coordinates of the final 

profile (inset in Figure 7) were delivered to NASA for subsequent testing in the transonic linear cascade as described in the 

following section.

3-D Blading Design/Efficiency-Speed Trend Analysis

Hub, midspan, and tip-sections were radially stacked on 

their center of gravity and were analyzed using 3-D RANS 

computations in ADPAC (Ref. 30) as an isolated bladerow. 

The computations were used to analyze the impact of altitude, 

turbine pressure ratio, and shaft-speed change. Consistent with 

the findings illustrated in Figure 6, Rotor 2 efficiency was 

found to increase by two points in moving from the cruise to 

takeoff operating condition (Ref. 11).

Experimental Testing of

Incidence-Tolerant Blading

The GRC Transonic Turbine Blade Cascade (see Figure 8) 

has been used to obtain detailed flow and heat transfer data on 

a wide variety of turbine blade geometries. The tunnel operational capabilities (see Figure 9) enable testing at relevant LPT 

relative Mach numbers (0.6 < M2 < 1) and inlet turbulence 

intensities (0.3 percent < Tu < 15 percent), over an order-ofmagnitude range of Re, and wide range of inlet flow (and 

incidence) angle settings (–17° < β1 < +70°). The cascade has 

been used previously to acquire heat transfer data sets

(Ref. 31). In the present effort, heat transfer rates have proven 

to be an effective marker for transitional flows of importance

to VSPT blading.

E3 Tip-Section Tests

Aerodynamic tests of a 2-D tip section of a NASA/GE 

Energy Efficient Engine (E3

) blade (Ref. 32) were performed 

as an initial step to get detailed code and model assessment 

data relevant to a VSPT. Midspan data were obtained 

8.6 percent cx downstream of the blade trailing edge over three 

blade passages and include total pressure, Mach number, and 

flow angle. A five-hole pneumatic probe was used, following 

the methodology described elsewhere (Ref. 33). Sample total 

pressure data are shown in Figure 10 at +4.1° and +19.1°

incidence angles. The incidence angle is defined based on a 

design-intent β1 = +29.7°. The design inlet Reynolds number 

is Recx,1 = 6.83 × 105 and the design pressure ratio is 1.441 

giving an isentropic exit Mach number of M2,i = 0.74. 

Other conditions were chosen to obtain as low a Reynolds 

number as possible within the operating envelope of the 

facility (Figure 9). The left side of each peak in Figure 10

corresponds to the pressure side of the wake while the right 

sides correspond to the suction sides. For i = +4.1°, the wake 

widths and depths grow with decreasing Reynolds number but 

all share the same linear suction surface profile. At a more 

highly loaded condition of i = +19°, the suction sides show 

significant thickening at lower Reynolds numbers, indicative 

of rapid boundary layer growth brought about by incipient 

separation. These effects are anticipated for VSPT blading as 

well, making this an appropriate test case for predictive model 

validation.
NASA/TM—2012-217605 9

Figure 8.—Schematic drawing of the GRC transonic linear cascade, recently 

modified for testing incidence-tolerant blading for VSPT.

Figure 9.—GRC transonic linear cascade operating map defined by 

flow boundaries, showing VSPT and previous heat transfer and 

aerodynamic test points.
NASA/TM—2012-217605 10

(a) +4.1° incidence

(b) +19.1° incidence

Figure 10.—Aero survey data acquired in the NASA transonic 

linear cascade at midspan of E3 blading at design pressure 

ratio and varying Reynolds number, for two inlet flow angles.

Incidence-Tolerant VSPT Blade Tests

During the past year, the cascade facility (NASA CW-22) 

was modified significantly to extend the range of flow angle 

setting so as to accommodate high negative incidence testing 

of relevance to VSPT blading. The tunnel was re-baselined 

using the E3 blading by repeating the test points at various 

inlet flow angles, including the positive incidence angles 

described in the previous section. 

The midspan section of the incidence-tolerant VSPT rotor 

blading (Figure 7 inset) developed under the RTAPS contract 

with RR-NAT (Ref. 10) is currently being tested in the 

modified facility. Loss buckets of the type shown in Figure 7

are being generated for a parametric range of Reynolds 

numbers. Wake data were acquired 7.0 percent cx downstream 

of the VSPT blade trailing edge using a 5-hole probe. A 3-hole 

boundary-layer probe was used in the region below 4.2

percent of span from the endwall.

(a) Recx,2 = 0.212×106, M2,i = 0.35

(b) Recx,2 = 2.120×106, M2,i = 0.72

Figure 11.—Aero survey data acquired in the NASA transonic 

linear cascade at midspan of the RR-NAT designed VSPT 

blading at various Reynolds numbers, pressure ratios, and 

inlet flow angles.

Total pressure coefficient data are shown in Figure 11 at the 

minimum and maximum Reynolds number conditions and 

over a wide range of inlet (and incidence) angles; β1 = +40.0°

(i = +5.8°, cruise), β1 = –2.5° (i = –36.7°, takeoff), and 

β1 = –11.8° (i = –46.0°, max. incidence). The pressure surface 

is on the left side of the wakes while the suction surface is on 

the right. The data clearly show the effects of angle and flow 

condition on the pressure and suction surface wakes, again 

making this an appropriate test case for predictive model 

validation. Midspan data are being acquired at other inlet 

angles and flow conditions in order to give a more complete 

assessment of the losses. Full-span data (endwall to midspan) 

were also acquired at the lowest Reynolds numbers achievable 

at the design blade pressure ratio (M2 = 0.72) and at incidenceangles corresponding to the cruise (i = +5.8° design-point) and 

takeoff incidence levels (i = –36.7°).

 

 
NASA/TM—2012-217605 11

The incidence-tolerant VSPT blades were instrumented with 

static pressure taps to provide blade loading data to be correlated with the loss measurements. The majority of the pressure 

taps were located at midspan, but taps were also located at 10, 

15, and 30 percent span. The center measurement blade was 

instrumented most densely, and the suction and pressure 

surfaces of adjacent blades were instrumented to verify 

periodicity. Sample blade loading data are shown in Figure 12

for two flow conditions. The loading diagram in Figure 12(b)

shows a large cove separation region on the front portion of 

the pressure surface which leads to the thickening of the 

pressure surface wake shown in the wake measurements of 

Figure 11(a) at β1 = –11.8°. Conversely, the loading diagram 

in Figure 12(a) indicates attached flow everywhere, leading to 

the relatively narrow wakes shown in Figure 11(b). The 

loading measurements, when analyzed with the wake measurements, provide for a more complete interpretation of the 

effects of incidence and flow conditions.

(a) β1 = +40.0° (i = +5.8°), Recx,2 = 2.120×106, M2,i = 0.72

(b) β1 = –11.8° (i = –46.0°), Recx,2 = 0.212×106, M2,i = 0.35

Figure 12.—Static pressure blade loading from the incidencetolerant VSPT blading at various Reynolds numbers, 

pressure ratios, and inlet flow angles.

Computational Methods

The unit Reynolds numbers of the LCTR VSPT aft-stages 

are near 50k/in. The low Reynolds number operation admits 

transitional flow fields with significant laminar regions which 

are subject to separation induced transition to turbulent flow. 

The transitional flow is sensitive to the deterministic buffeting 

of wakes from upstream blade rows (Refs. 5, 26, and 27). 

Thus, in addition to the difficulty of predicting 3-D transitional flow fields computationally due to turbulence/transition 

sub-model limitations, the impact of unsteadiness must be 

included in accurate predictions of turbine performance. The 

large incidence swings of the VSPT add an additional complication and degree of complexity to the problem as well. An 

internal effort (Ref. 34) to down-select and validate an 

accurate turbulence/transition model for RANS computations,

and for subsequent implementation in an unsteady RANS 

(URANS) solver for multistage turbine calculations, is 

reported in this section.

Turbulence Model for Transitional Flow Fields

Following a year-long investigation of various turbulence 

modeling approaches suitable for LPT flow fields, the Walters 

and Leylek (Refs. 35 and 36) model was implemented in the

NASA RANS code GlennHT (Ref. 37). The Walters-Leylek 

(W-L) model is a three-equation (k-ω-kl), model which 

accounts for the laminar kinetic energy, kl, portion of the total 

intensity of unsteadiness and its natural growth, admitting a 

mechanism for natural and bypass transition in the RANS submodel. The model implementation and verification is 

described in detail elsewhere (Ref. 38). The new code capability was assessed by comparisons of 2-D heat transfer for 

textbook (Ref. 39) flat plate flow fields and with transition 

heat transfer data sets obtained from experiments in the NASA 

transonic linear cascade (Ref. 31). The wall heat transfer data 

sets were used because heat transfer rates (Nu) provide a 

strong blade-surface marker for transitional flows. Upon 

ensuring the satisfactory performance of the transition model 

for test cases, the flow solver with the improved transition 

model is used to compute the losses.

Steady-State Heat Transfer Assessments

The results of 2-D and 3-D computations of blade heat 

transfer and comparison with experimental data for the GE2 

blade pack (Ref. 31) are shown in Figure 13 to Figure 15

(reproduced from Ref. 38). The GE2 blade is an extruded 2-D 

section of a high-turning (127°) power-turbine blade. The data 

were obtained in the GRC transonic tunnel at two Recx,2 while 

the pressure ratio (and M2) was held constant. 

 
NASA/TM—2012-217605 12

Figure 13.—Two-dimensional computation of Nusselt number against data of Giel et al. (Ref. 31), at 

Recx,2 = 375k.

Figure 14.—Two-dimensional computation of Nusselt number against data of Giel et al. (Ref. 31), at 

Recx,2 = 621k.

Figure 15.—Three-dimensional computation of Nusselt number and comparison with data from Giel 

et al. (Ref. 31), Recx,2 = 375k.
NASA/TM—2012-217605 13

Figure 16.—Comparison of measured1 (symbols) and computed (curves) totalpressure coefficients at inlet flow angles, α (in deg.), and corresponding incidence 

angles (parenthesis, in deg.) through VSPT-relevant range of –47° < i-ides < +20°

for E3 tip-section blading (Recx,1 = 83k).

The computed heat transfer rates of the 2-D RANS/W-L 

computation are in excellent agreement experimental data on 

the suction side, but heat transfer rates on the pressure-side are

under-predicted. The root-cause for the disagreement on the 

pressure side has not yet been determined. It is noted that the 

cascade flow field is highly three-dimensional: the inlet 

boundary layer on each endwall is approximately 15 percent

of span and the aspect ratio of the blade pack is less than 

unity; combined with the high turning of the GE2 cascade, the 

degree of agreement with 2-D calculations was unexpected.

The 3-D computations for the same cascade are provided in 

Figure 15. A comparison of the computational and experimental data indicates that the impact of the strong secondaryflow fields and associated high aerodynamic blockage levels 

may be overestimated by the simulation. The location of 

transition onset is, nonetheless, well predicted at the three 

spanwise measurement locations.1

Computations of Wake Profiles and Loss in 

Steady-State Cascade

The agreement for transition between CFD predictions and 

data from the GE2 cascade test is considered to be quite good. 

The GlennHT code with the W-L turbulence/transition model 

was further exercised to compute loss as a function of inci-

 1

The data were obtained by A. McVetta and P. Giel, 2011; to 

be published in AIAA Joint Propulsion Conference, July 2012.

dence for a cascade composed of E3 tip-section blading. Data 

were again obtained in the GRC transonic linear cascade as 

described in the previous section. Experimental data from exit 

surveys with 3-hole p0 probes were used for verification over a 

wide (detail) range of incidence angles, exit Mach numbers, 

and Reynolds numbers. Results for M2 = 0.34 and Recx,1 = 

83k are provided here. For the computations, Tu = 1 percent

and the inlet boundary layer thickness is 49 percent of halfspan.

A comparison of computed and measured midspan totalpressure coefficients is provided in Figure 16. The totalpressures are acquired at an axial station 8.6 percent cx

downstream of the trailing edge of the cascade. The incidence 

angle is defined based on a design-intent β1 = +29.7°. Integrated wake deficits (loss) are highest at the positive incidence 

angles. Interestingly, the agreement between CFD and data is 

best at extreme positive (most loaded) and negative (unloaded) 

flow angles; the disagreement in wake depths at the neardesign inlet angle (β1 = +33.8°, i-ides = +4.1) is substantial and 

unexplained at this time.

Cross-passage contours of Cptot for incidence angles of approximately –65°, –15°, +10° and +30° are shown in Figure 17

and iso-surfaces of Cptot are provide in Figure 18. The results in 

Figure 17 and Figure 18 were carried out at M2 = 0.35 and Recx,1

= 100 k. Note that only hub to the midspan symmetry plane is 

shown. The total-pressure deficits can be identified for each 

case. The highest loss regions are associated with the vortex 

cores. This is especially true at the higher positive incidence
NASA/TM—2012-217605 14

Figure 17.—Total-pressure loss contours (halfspan) of E3 tip-section blading from (clockwise 

from top-left) for approximate incidences of 

(a) –65°, (b) –15°, (c) +10°, and (d) +30° (from 

Ref. 38). Note contour scale change.

Figure 18.—Iso-surfaces of computed loss coefficient of E3 tipsection blading over half-span for approximate incidence 

angles of a.) –65°, b.) –25°, c.) +10°, and d.) +30° (from 

Ref. 38).

angles at which loading and secondary flows are highest. 

Corresponding flow field structures are provided in Figure 18

in terms of iso-surfaces of total-pressure coefficient. The 

patterns of Cptot follow, for the most part, the vortical structures highlighted by the iso-surfaces. The contour levels are 

chosen to illuminate the flow structures, and the iso-surfaces 

are colored by density to allow better visualization.

For Case (a) with extreme negative incidence, the largest 

loss near the endwall is due to the interaction of endwall 

boundary layer and the merged horseshoe/passage vortex 

system, which stays near the endwall due to the unloading of 

the blades (Figure 17(a)). The mid-passage increase in loss for 

Case (a) is due to large flow separation on the pressure side 

(cove separation) which starts up as a spanwise vortex but 

turns into a streamwise vortex near midspan. For the other 

three cases, the midspan loss may be attributed to the suction 

side leg of horseshoe vortex. High loss is reflected as a bulge 

in the loss contours in the wake of Case (b). This is highlighted by the better formed local maxima for Cases (c) and (d), 

and is due to the combined effect of horseshoe and passage 

vortices. As the loading increases with the incidence, the 

magnitude and extent of the high loss region increases. 

Integrated Results 

Using computed wake profiles at spanwise sections between 

endwall and midspan, loss coefficients based on area-averaged 

total-pressures were calculated (Figure 19). The resulting loss 

bucket for the E3 tip-section blading has extensive (70°) range 

of incidence at design-point loss levels.

Multistage Turbine Simulations

With intent to improve in-house multistage turbine simulation capability, a NASA supported unsteady turbomachinery 

code, TURBO (Ref. 40), is currently under further development and evaluation in support of multistage simulation

capability for VSPTs with transitional flow. The code improvements were to be implemented in a two-part approach. 

Firstly, TURBO was to be run against documented multistage 

turbine geometry so that the mechanics of single-passage per 

bladerow, phase-lag simulations could be exercised and 

refined (Ref. 41). Secondly, the W-L turbulence/transition 

sub-model (Refs. 35 and 36) described above was to be 

implemented in the TURBO code. The model improvements 

are described in this section.

Application to Low-Speed Rotating Rig

TURBO (Ref. 40) is a 3-D unsteady RANS (URANS) multi-blade-row, multi-block aero code; the code has been applied 

to-date primarily to compression systems. As reported earlier

(Ref. 41), an assessment of code performance was conducted

using the United Technologies Corporation (UTC) large scale 

rotating rig (LSRR) in 1.5-stage (S1/R1/S2) turbine configuration (Ref. 42). The multistage URANS computations were

performed in single-passage per blade-row, phase-lag mode. 

The simulation details and related analysis are documented

elsewhere (Ref. 41). Example time-averaged results are 

provided in Figure 20, where measured and computed timeaveraged loading diagrams on Rotor 1 and Stator 2 of the 

LSRR are compared and were found to be in acceptable 

agreement (Ref. 41). Computed results for two flow 

a. b.

d. c.

c. d. 

a. b. 
NASA/TM—2012-217605 15

Figure 19.—Loss coefficient based on area-averaged total-pressures of 3-D loss 

integration of flow fields (Figure 17) as function of incidence angle for extruded E3 tipsection blading tested in NASA transonic linear cascade at Recx,2 = 100k.

(a) Rotor 1 (b) Stator 2

Figure 20.—Example results from multistage TURBO computation (Ref. 41), showing comparison of computed and 

measured time-averaged static pressure on (a) Rotor 1 and (b) Stator 2 of Low Speed Rotating Rig (Ref. 42).

coefficients ( φ = ux /U ), 0.76 and 0.77, were provided to 

illustrate the sensitivity of the computed loading diagrams to 

flow coefficient variation.

The main conclusion from the multistage URANS investigation is that a number of the standard practices used when 

simulating fans and compressors with TURBO have to be 

modified for the turbine application. Firstly, clearance gaps 

will need to be gridded and clearance flows computed; the 

thickness of the turbine blades renders the tip-clearance 

models arbitrary and insufficiently accurate. The validity of 

wall-functions used with the k-ε model in TURBO is questionable as the mesh is refined to compute the clearance flow. 

Secondly, because of the large blade turning angles, a mesh 

topology such as an O-H mixed type or a multi-zone type, as 

opposed to the H or C type, is more appropriate for turbine 

simulations. Thirdly, the use of wall functions in the current kε turbulence model is not expected to do well in low-Re

transitional flow.

0 0.2 0.4 0.6 0.8 1.0

x/cx

Experimental data at phi = 0.78 +/- 1%

Cp

2

3

5

10

suction side

pressure side

1

4

6

7

8

9

 

0 0.2 0.4 0.6 0.8 1.0

x/cx

Experimental data at phi = 0.78 +/- 1%

Cp

4

6

8

10

12

suction side

pressure side
NASA/TM—2012-217605 16

Implementation of Walters-Leylek Model in TURBO

The 3-equation W-L turbulence/transitional-flow model

(Refs. 35 and 36) described above was implemented in 

TURBO and is being evaluated currently against heat transfer 

data from the linear cascade experiments with GE2 blading

(Ref. 31) in which the Reynolds numbers are sufficiently low 

to admit transitional flow. As in the GlennHT implementation

described in the previous section, integration to the wall, 

without the use of wall functions permitted (and warranted) 

with the W-L model, and therefore the y+ ~ 1 or less at the 

first point from the wall in future URANS computations.

Rotordynamics

At the initiation of the VSPT research effort, there existed 

uncertainty about the feasibility of the concept related to the 

rotordynamic behavior of a multiple-rotor engine operating 

over a wide speed range (54 to 100 percent of takeoff speed). 

The source of uncertainty was related to the ability to manage 

rotor vibration response at operating speeds, and while 

traversing critical speeds during ramp-up/ramp-down. A 

rotordynamic feasibility analysis was launched in order to 

assess the impact of the variable speed concept on an engine in 

the 7500 shp-class suited for the NASA LCTR. The study

(Ref. 6), summarized in this section, was focused on rotor 

natural frequencies (critical speeds), stability, and steady-state 

unbalance response of the concept engine with intent to 

determine the plausibility of the VSPT concept in this size 

class.

NASA Assessment of Rotordynamics

A rotordynamic model of the notional engine was developed using geometry from a concept engine (Ref. 43) developed for NASA’s Large Civil Tilt Rotor (LCTR2) concept 

vehicle. The front-drive engine model consists of three rotors: 

a high-speed (33,000 rpm) high-pressure compressor rotor 

(HP), a moderate-speed (27,500 rpm) low-pressure compressor rotor (LP), and a lower speed (8,100 to 15,000 rpm) VSPT 

rotor. The size and mass of the various stages, and the overall 

layout of the engine, were taken from the system study 

analysis for the engine (Ref. 43). 

The rotordynamic model was developed in a commercially 

available rotordynamics software package that uses a finite 

element-based, 1-D (axisymmetric), lumped-parameter 

method to simulate the geometry. A schematic of the three 

rotor model of the notional engine is provided in Figure 21. A 

detailed analysis of the feasibility of the concept engine is 

provided elsewhere (Ref. 6), and a summary of the analyses 

and findings is given here.

Critical Speed Analysis

An important aspect of any rotating system is the existences 

of natural frequencies of the rotating structure that can be 

excited by various sources and potentially lead to failure. 

Traditionally, aero engines have been designed with relatively 

fixed-speed power turbine rotors during nominal operation. 

These rotor systems need to accelerate up to nominal speed, so 

clearly natural frequencies must be understood; nonetheless, 

the narrow operational speed range of the typical rotorcraft PT 

shaft eases the difficulty with which critical speeds can be 

avoided during steady-state operation. The addition of variable-speed capability in the power turbine complicates the 

avoidance of natural frequencies in the operating speed range. 

The larger the operating speed range, the more rotor modes are 

spanned and may need to be avoided. 

For the current VSPT research effort, the rotordynamic 

model (Figure 21) was exercised to iterate towards a design 

that eliminated critical speeds from the operating speed range 

of the PT rotor. In addition, the critical speeds of the LP and 

HP rotors were calculated to quantify the margin between the 

critical speed frequencies and the running speed of the 

respective rotor. No iteration was done on the LP or HP rotors, 

but the response of the HP and LP rotors could be refined as 

warranted during detailed design.

A Campbell Diagram for the rotordynamics of the three 

spools of the notional LCTR engine is provided in Figure 22. 

The natural frequencies/modes of the three rotors are plotted

as a function of rotor speed. The first-engine-order (1EO) line 

is denoted 1X and represents a synchronous excitation 

(usually unbalance). The diagram can be used to identify the 

speeds at which an excitation force—in this case the synchronous unbalance force—coexists with the natural frequencies of 

the various rotors. The intersections of the 1X line with the 

natural frequencies define the critical speeds or “crossings.” 

To ensure robust mechanical design, the 1X crossings within 

the VSPT shaft operating range were avoided; as indicated in 

Figure 22, there are no 1EO intersections in the power turbine 

operating range, or near the LP rotor and HP rotor design 

points. 

In order to achieve the desired critical speed behavior, several iterations on the PT rotor design were required. Additional optimization is possible, but may require modifications to 

the geometry of the LP and HP rotors, and is left to detailed 

design. In more detailed efforts, higher engine-order crossings 

may need to be considered as well, as well variations in HPand LP-shaft speeds over the mission. Nonetheless, feasibility 

from the standpoint of critical speeds was demonstrated by the

preliminary level analysis (Ref. 6).
NASA/TM—2012-217605 17

Figure 21.—Rotordynamic model of 3-spool, front drive concept engine for 

NASA LCTR application.

Figure 22.—Campbell Diagram for three rotors of the LCTR concept engine.
NASA/TM—2012-217605 18

Stability Analysis

Stability is an important consideration in aero-engines in 

which relatively lightweight rotors, spin at relatively high 

speeds and power, with relatively low damping. This combination of factors can result in an unsafe system susceptible to 

instability within the operating range. For the conceptual 

LCTR2 engine, stability could not be analyzed rigorously 

because the details of the bearings, bearing supports, and 

housing are unknown; nonetheless, to estimate susceptibility of 

the design to instability, a simplified stability assessment was 

undertaken. An artificial amount of damping is added to each 

bearing support. The amount of damping was incrementally 

increased at each bearing location until all three rotors of the 

engine were stable up to their respective maximum operating 

speeds. In this manner, rather than assessing whether the system 

was stable with a known amount of damping, the amount of 

damping needed to achieve stability was determined. Since the 

required levels of damping were found to be consistent with 

those of similar applications, stability of the conceptual design 

was deemed achievable. A detailed description of the stability 

analysis is provided in Reference 6.

Unbalance Response

Unbalance response sets the rotor balance requirements to 

ensure that rotor system displacements remain within allowable limits, both at full-speed steady-state operation and during 

resonance crossings during speed-up and speed-down events. 

While passing through resonances, the amplitude of motion at 

the rotor components must not exceed the transient clearances 

or rubs will occur. Additionally, industry standards for

acceptable vibration amplitudes at operating speed must be 

met, and bearing reaction forces while traversing resonances 

must be assessed.

To assess the viability of the current conceptual design, a 

typical new-machine unbalance specification was applied to 

the rotordynamic model, and the response was calculated over 

the speed range for each rotor and compared to the accepted 

industry standard (Ref. 44). The resulting peak-to-peak 

displacement amplitudes (Xpk-pk) calculated for the PT rotor at 

low and high speeds, and the industry specification are 

provided in Table II. Since the unbalance response amplitude 

was found to be less than the standard allowable displacements, this design is considered to be viable from the standpoint of balance considerations. Details of the unbalance 

response analysis are provided in Reference 6.

TABLE II.—VIBRATION AMPLITUDE FOR 

VSPT OF LCTR CONCEPT ENGINE

Speed,

rpm

Xpk-pk at front 

bearing location,

mm

Xpk-pk at rear 

bearing location,

mm

Allowable

pk-pk amplitude,

mm

8100 5.08×10–4 8.89×10–3 2.11×10–2

15000 3.81×10–3 8.64×10–3 1.13×10–2

Industry Assessments of Rotordynamics of 

LCTR Engines

RR-NAT Assessment

In the NASA RTAPS study contract by RR-NAT (Ref. 11), 

a growth version of an existing front-drive two-spool 

AE1107C engine was modified to meet the LCTR requirements. A third stage was added to the VSPT to extend the 

operational speed-range of the turbine. A revised model of the 

AE1107C was created for use in the rotordynamics analysis. 

Subsequently modifications were made to the VSPT concept 

engine, including PT shaft thickening and modifications to the 

aft PT bearing placement, and the shaft attachment (from PT 

disk 1 to 2) in an effort to modify the response and frequency 

placement of potential responsive modes. The predicted 

bearing loads for a 1-mm out-of-balance (imbalance response 

analysis) for the AE1107C baseline and the modified 3-stage 

VSPT variants, HP and PT rotors, were compared. The 

bearing loads of the two engine configurations were in close 

agreement; RR-NAT concluded that the modified AE1107C 

engine configuration with incorporated VSPT technology is 

rotordynamically viable (Ref. 11).

Williams International Assessment

An aft-drive engine configuration concept was proposed by 

WI for the LCTR mission. The VSPT shaft emanates from the 

aft of the engine, as in the P&WC PT6 turboprop engine, to 

drive a transmission of fixed reduction gear-ratio. An advantage of this approach is a shorter shaft, relative to shafts of 

the more common front-drive engine configurations, which 

could strongly reduce torque and rotordynamics issues. The 

configuration holds potential engine protection advantages as 

well. Disadvantages are associated with the need to turn the 

ambient intake and hot exhaust streams by 180°. The WI 

analysis was not carried out to a level of detail required to 

assess the placement of critical modes of the concept engine, 

however the implication is that the aft-drive configuration 

offers flexibility in terms of the shaft diameter selection and 

configuration to achieve required critical speed margins

(Ref. 9).

Summary of Rotordynamic Assessments

The conclusion from the NASA in-house and engine company studies is that rotordynamically viable mechanical 

designs for the VSPT of the LCTR class engine are achievable. The more common front-drive system of the 7500 shpclass engine will likely operate with a supercritical rotor in 

which VSPT shaft criticals are within the VSPT shaft-speed 

range. Related rotor-rotor interactions will need to be avoided 

by mode placement away from steady-state operating points,

and responses during transient crossings managed by incorporation of suitable damping levels and shaft thicknesses. The 
NASA/TM—2012-217605 19

aft-drive configuration may admit a subcritical VSPT-shaft 

solution in which all VSPT-shaft modes are above the engine 

running range.2 If a subcritical configuration (short, large 

diameter shaft) were achievable, and the aft-drive configuration proves compatible with aerodynamic, heat load, and 

packaging constraints, the approach would eliminate VSPT 

criticals and interactions with other rotor modes within the 

VSPT running range.

VSPT Component Testing

Potential at GRC

The results from an assessment of potential paths to component VSPT component in-house testing at GRC, informed by 

findings from the VSPT study contracts with RR-NAT

(Ref. 11) and WI (Ref. 9), are summarized in this section.

Two in-house approaches to VSPT component testing were 

considered (Ref. 45): i.) experimental testing would be 

conducted in an existing power turbine of a T700 engine 

located in the GRC Engine Components Research Laboratory 

(ECRL); or, ii.) a new VSPT component rig would be installed 

in the new single-spool turbine test facility (W-6) at GRC. The 

objective of the assessment study was to determine the level of 

compatibility that exists between each research facility and the

 2

Note that the analysis and discussion provided in the Williams Int. study contract neither supported nor refuted the 

supposition that a subcritical VSPT shaft may be enabled by 

the aft-drive configuration.

VSPT component research needs. Preliminary research 

requirements were defined and the experimental assets 

available in each facility were examined to determine which 

approach would provide a greater return on the time and funds 

to be invested in VSPT experimentation.

Instrumented T700 PT in ECRL

The Engine Component Research Laboratory (ECRL) of 

GRC is currently configured as a sea-level turboshaft engine 

test stand. A recently completed assessment has shown that 

the T700-700 power turbine could be operated with research 

instrumentation (Figure 23 and Figure 24) in ECRL at the 

normalized speeds and loads that would simulate the requirements of a notional LCTR mission profile. While provisions 

for altitude testing exist, the low Reynolds numbers of cruise

altitudes could not be matched without a facility upgrade. A 

rotordynamic model of the T700 power-turbine shaft and 

support system was used to identify critical speeds of potentially responsive shaft modes. These analyses, along with 

previously acquired engine vibration data, supported the 

conclusion that the PT speed could be operated with acceptable shaft response at the design speed and half-speed conditions. The need to modify the engine control system was 

identified as well.

Figure 23.—T700 cross-section, showing proposed positions of rating rakes and 

static pressure measurements. 
NASA/TM—2012-217605 20

Figure 24.—T700 flat layout showing position of total-temperature and pressure elements and swirl 

angle capability of downstream rake (Ref. 45).

An instrumentation plan was developed to provide for 

measurements within the power turbine that would include 

overall performance measurements (T0/p0 rakes at S4.5 and 

S5.0 of Figure 23) and inter-stage static pressure measurements, as well as provisions for radial surveys of total pressure, total temperature and flow angle (5-hole probe) at the 

power turbine exit (Station 5.0 of Figure 24). Testing could 

begin after routine maintenance and safety upgrades to the 

facility were completed and the aforementioned instrumentation was installed in the T700 power turbine module. These 

tasks could be performed concurrently and were estimated to 

take roughly 1.5 years to complete at a relatively low initial 

cost. As the existing power turbine would be used, this 

approach would provide data to calibrate analytical and design 

methods but not to evaluate a specific VSPT design.

VSPT Component Rig in Single-Spool Turbine 

Facility

The GRC W-6 Single-Spool Turbine Facility (Figure 25) 

has heated (warm) inlet conditions (940 °F max) and massflow-rate (27 lbm/s max) and pressure (50 psia max.) range 

suitable for testing components that are full-scale, or are 

scaled-up, as favorable for detailed measurements of flow 

profiles and performance. The GRC lab altitude exhaust 

(2 psia) system would accommodate mission-relevant variations in Reynolds number within current facility capabilities. 

Heated secondary air supply systems are available as well. 

Intake flow rates are measured with a calibrated venturi and 

torque is measured using a stand-alone torque meter. Power 

transmission and absorption is through a reduction gear-box to 

a synchronous generator (Figure 25) with tight speed control 

vis-à-vis the lab-wide variable frequency system. The W-6 

facility will not become available until 2014 due to completion 

schedule and initial check-out and baseline testing commitments.

As part of the VSPT study contracts, Rolls-Royce NAT and 

Williams Int. each developed a plan for a cost-effective test 

approach to conduct component-level research for the LCTRrelevant VSPT. The approach included a research and instrumentation plan, test plan, schedule, and cost estimate. In both 

studies, the proposed approach was to test a multistage (3.5- or 

4-stage) VSPT at scale in the NASA W-6 facility. The

experiments would accommodate a multistage VSPT designed 

to match LCTR engine Mach number, pressure and temperature ratios, and flight Reynolds number requirements. A

notional 4-stage test article installation in W-6 and performance rating instrumentation layout from the WI VSPT study

contract (Ref. 9) is provided in Figure 26.

Assessment Results

Both the ECRL/T700 engine facility and the W-6 turbine 

facility were found to be viable options for VSPT experiments. The initial time and funding requirements to prepare an 

experiment in ECRL were estimated to be less than for the W6 facility, provided the existing T700 power turbine was 

retained. Using information documented in the VSPT study 

contracts(Refs. 9 and 11), the initial cost of the W-6 component approach was estimated to be three to four times higher 

than the ECRL/T700 approach (using the existing T700 power 

turbine), and the estimated weekly facility operating costs

 
NASA/TM—2012-217605 21

Figure 25.—Schematic diagram of the GRC W-6 Single Spool Turbine Facility.

Figure 26.—Williams International 4-stage LCTR VSPT component in GRC W-6 single-spool 

facility, showing notional rating instrumentation for VSPT turbine-performance mapping 

(Ref. 9).

air supply;

40 psig Combustion Air

(clean, dry, ambient temperature)

gearbox

sync.

machine

Turbine Test Article

work platform for

secondary air circuits

air heaters

flow measurement

venturi

to Altitude

Exhaust

at 26” Hg

flow

flow

Temperature 

Probes

Tt/Pt 

Rakes 

Pressure taps on all stages NASA Drive System Surrounding Vane Airfoils

On Inner and Outer Flowpaths

Clearance Rub Pins

Mate Rig to

Drive System

Mate Facility 

Collector

Existing Facility 

Instrumentation –

speed, torque, exit 

gas conditions, 

health.
NASA/TM—2012-217605 22

were roughly equal. The T700 engine approach would 

constitute a first step toward verification of the operational 

capability of the engine and power-absorber over the 50

percent speed range and of the ability to acquire researchquality data from the instrumented power turbine module 

when integrated in the engine. The PT module would need to 

be modified/replaced with a compatible VSPT module in a 

subsequent step. The W-6 facility was found to provide better 

access for research instrumentation and would permit researchers to evaluate a full-scale power turbine to support the 

specific aerodynamic research objectives. The W-6 facility is 

currently the preferred candidate for future VSPT componentlevel aerodynamic research experiments.

Conclusions

The variable-speed power turbine (VSPT) research effort at 

GRC was reviewed. In-house and engine-company conceptual 

design studies have indicated that the VSPT for the NASA 

LCTR mission will be a 3.5- to 4.5-stage, fixed geometry 

turbine, dependent in part on the final power requirements and 

engine power-to-weight-flow levels. Nominal turbine speed 

operation of 8100 rpm < N < 15000 rpm would correspond to 

the 54 percent < N* < 100 percent speed requirement. On- and 

off-design steady and unsteady 3-D multistage computations 

completed to-date corroborate meanline analyses in terms of 

expected increase in turbine efficiency between the cruise (54

percent N*

) to takeoff (100 percent N*) flight points. The offdesign correlations of the meanline tools are critical for such 

concept studies and improvements at the extreme negative 

incidence levels are deemed warranted. The VSPT performance maps and insights generated from the VSPT concept 

design studies and analyses are informing system studies 

intended to refine the NASA LCTR mission.

Initial experimental results for incidence-tolerant blading 

tested over a range of VSPT relevant incidence angles were 

generated in the recently modified NASA transonic linear 

cascade. The presented results demonstrate that tunnel 

capability spans the wide range of Reynolds numbers needed 

to test blading from fully turbulent to transitional flow 

conditions.

Implementation of the Walters-Leylek turbulence model 

into in-house 3-D RANS and URANS solvers was described. 

The W-L model is found to predict accurately the critical 

suction-side transition, based on comparisons with steady-state 

heat transfer data previously obtained in the transonic linear 

cascade. Validation using unsteady data from a high-speed 

multistage component rig remains for future work.

Conceptual design level rotordynamics analyses, conducted 

in-house and through the VSPT study contracts with industry, 

indicate that rotor-dynamically viable engine concepts are 

obtainable for the 7500 shp-class engine with VSPT.

Finally, two paths to engine testing using internal assets 

were explored and found to be viable. The preferred path to 

component test is the use of a turbine test facility with mission 

Reynolds number test capability, such as the NASA W-6 

Single-Spool Turbine Facility.

References

1. Johnson, W., Yamauchi, G. K., and Watts, M. E., 

“NASA Heavy Lift Rotorcraft Systems Investigation,” 

NASA/TP—2005-213467, Sep. 2005.

2. Acree, C. W., Hyeonsoo, Y., and Sinsay, J. D., “Performance Optimization of the NASA Large Civil Tiltrotor,” 

Proc. Int. Powered Lift Conf., Jul. 2008.

3. Wilkerson, J. B. and Smith, R. L., “Aircraft System 

Analysis of Technology Benefits to Civil Transport Rotorcraft,” NASA/CR–2009-214594, Jun. 2009.

4. Welch, G. E., “Assessment of Aerodynamic Challenges 

of a Variable-Speed Power Turbine for Large Civil TiltRotor Application,” Proc. AHS Int. 66th Ann. Forum, 

May, 2010; also NASA/TM—2010-216758, Aug. 2010.

5. Hourmouziadis, J., “Aerodynamic Design of Low Pressure Turbines,” in Blading Design for Axial Turbomachines, AGARD-LS-167, 1989, pp. 8-1 to 8-40.

6. Howard, S. A., “Rotordynamic Feasibility of a Conceptual Variable-Speed Power Turbine Propulsion System 

for Large Civil Tilt-Rotor Applications,” Proc. AHS Int. 

68th Ann. Forum, TX, May 2012; also NASA/TM—

2012-217134, May 2012.

7. Stevens, M. A., Handschuh, R. F., Lewicki, D. G., “Variable/Multispeed Rotorcraft Drive System Concepts,” 

NASA/TM—2009-215456; also ARL-TR-4728, Mar.

2009.

8. Litt, J. S., Edwards, J. M., DeCastro, J. A., “A Sequential 

Shifting Algorithm for Variable Rotor Speed Control,” 

NASA/TM—2007-214842, June, 2007; also, ARL–TR–

4086, June 2007. 

9. Suchezky, M., “Variable-speed power turbine for the 

Large Civil Tilt Rotor,” NASA/CR—2012-217424, Feb.

2012.

10. Ford, A., Turner, E., Bloxham, M., Gegg, S., King, B., 

Harris, C., Bell, M., Eames, D., “Variable-Speed PowerTurbine Component Research Plan,” NASA/CR—2012-

217423, Feb. 2012.

11. Ford, A., Bloxham, M., Turner, E., Clemens, E., Gegg, 

S., “Design Optimization of Incidence-Tolerant Blading 

Relevant to Large Civil Tilt-Rotor Power Turbine Applications,” NASA/CR—2011-217016, 2012.

12. Snyder, C. A. and Thurman, D. R., “Gas Turbine Characteristics for a Large Civil Tilt-Rotor (LCTR),” Proc. 

AHS Int. 65th Ann. Forum, May 2009; also NASA/TM—

2010-216089, Feb, 2010.

13. Welch, G. E., “Computational Assessment of the Aerodynamic Performance of a Variable-Speed Power Turbine for Large Civil Tilt-Rotor Application,” Proc. AHS 

Forum 67, May 2011; also NASA/TM—2011-217124, 

Nov. 2011.

14. Clark, J. P., Koch, P. J., Ooten, M. K., Johnson, J. J., 

Dagg, J., McQuilling, M. W., Huber, F, and Johnson, P. 
NASA/TM—2012-217605 23

D., “Design of Turbine Components to Answer Research 

Questions in Unsteady Aerodynamics and Heat Transfer,” AFRL-RZ-WP-TR-2009-2180, Sep. 2009.

15. Gier, J., Franke, M., Hübner, N. and Shröder, T., “Designing LP Turbines for Optimized Airfoil Lift,” J. Turbomachinery, 132, Jul. 2010, pp. 031008-1 to -12.

16. D′Angelo, M., “Wide Speed Range Turboshaft Study,” 

NASA/CR—1995-198380, Aug. 1995.

17. Smith, S. F., “A Simple Correlation of Turbine Efficiency,” J. Royal Aero. Soc., 69, 1965, pp. 467-470.

18. Oates, G. C, Aerothermodynamics of Aircraft Engine 

Components, AIAA Education Series, 1985.

19. Ainley, D. G. and Mathieson, G. C. R., “A Method of 

Performance Estimation for Axial-Flow Turbines,” Aeronautical Research Council (ARC), R&M 2974, 1957.

20. Dunham, J. and Came, P. M., “Improvements to the Ainley-Mathieson Method of Turbine Performance Prediction,” J. Eng. for Power, Jul. 1970, pp. 252-256.

21. Kacker S. C., and Okapuu. U., “A Mean Line Prediction 

Method for Axial Flow Turbine Efficiency,” J. Eng. 

Power, 104, Jan. 1982, pp. 111-119.

22. Chima, R. V., “Calculation of Multistage Turbomachinery Using Steady Characteristic Boundary Conditions,”

AIAA-98-0968, Jan. 1998; also NASA/TM—1998–

206613, Jan. 1998.

23. Welch, G. E., “Computational Study of the Impact of 

Unsteadiness on the Aerodynamic Performance of a Variable-Speed Power Turbine,” AIAA-2012-937, Jan.

2012; also NASA/TM—2012-217425, 2012.

24. Rose, M., Schüpbach, P., Mansour, M., “The Thermodynamics of Wake Blade Interaction in Axial Flow Turbines: Combined Experimental and Computational

Study,” ASME GT2009-59655, Jun. 2009.

25. Adamczyk, J. J., “Model Equation for Simulating Flows in 

Multistage Turbomachinery,” ASME GT1985-226, 1985.

26. Haselbach, F., Schiffer, H.-P., Horsman, M., Dressen, S., 

Harvey, N., and Read, S., “The Application of Ultra 

High Lift Blading in the BR715 LP Turbine,” J. Turbomachinery, 124, Jan. 2002, pp. 45-51.

27. Halstead, D. E., Wisler, D. C., Okishii, T. H., Walker, G. 

J., Hodson, H. P, Shin, H-W., “Boundary Layer Development in Axial Compressors and Turbines: Part 3 of 

4—LP Turbines,” J. Turbomachinery, 119, 1997, p. 225.

28. Ni, R-H, Humber, W., Fan, G., Johnson, P. D, Downs, J., 

Clark, J. P., Koch, P. J., “Conjugate Heat Transfer Analysis of a Film-Cooled Turbine Vane,” ASME GT2011-

45920, Jun. 2011.

29. Robuck, M., Wilkerson, J., Zhang, Y., Snyder, C., 

Nordstrom, C., “Design Study of Propulsion and Drive 

Systems for the Large Civil Tiltrotor (LCTR2) Rotorcraft,” Proc. AHS 67th Ann. Forum, May 2011.

30. Hall, E. J., and Delaney, R. A., “Investigation of Advanced Counter-rotation Blade Configuration Concepts 

for High Speed Turboprop Systems: Task V - ADPAC 

User's Manual,” NASA/CR—1995-195472, 1995.

31. Giel, P. W., Boyle, R. J., and Bunker, R. S., “Measurements and Predictions of Heat Transfer on Rotor Blades 

in a Transonic Turbine Cascade,” J. Turbomachinery, 

126, Jan. 2004, pp. 110-121.

32. Timko, L. P., “Energy Efficient Engine High Pressure 

Turbine Component Test Performance Report,” 

NASA/CR—1984-168289, 1984.

33. Giel, P. W., Thurman, D. R., Lopez, I., Boyle, R. J., 

VanFossen, G. J., Jett, T. A., Camperchioli, W. P., La, H., 

“Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade,” ASME 96-GT-113, Jun, 1996; also NASA/TM—1996-107388 and ARL-TR-1252.

34. Ameri, A. A., “Implementation of a Transition Model in 

a NASA Code and Validation using Heat Transfer Data 

on a Turbine Blade,” NASA/CR—2012-217436, 2012.

35. Walters, D. K., Leylek, J. H., “A New Model for Boundary Layer Transition Using a Single-Point RANS Approach,” ASME Journal of Turbomachinery, 126, (1), 

2004, p. 193.

36. Walters, D. K., Leylek, J. H., “Computational Fluid 

Dynamics Study of Wake-Induced Transition on a Compressor-Like Flat Plate,” ASME Journal of Turbomachinery, 127, (1), 2005, p. 52.

37. Steinthorsson, E., Liou, M.S., and Povinelli, L.A., “Development of an Explicit Multiblock/Multigrid Flow 

Solver for Viscous Flows in Complex Geometries,” 

AIAA-93- 2380, 1993.

38. Ameri, A. A., “Use of Transition Modeling to Enable the 

Computation of Losses for Variable-Speed Power Turbine,” ASME GT2012-69591, Jul. 2012; also 

NASA/CR—2012-217435, 2012.

39. Incropera, F. P. and DeWitt, D. P., Fundamentals of 

Heat Transfer, Wiley, NY, 1981.

40. Chen, J. P. and Briley, W. R., “A Parallel Flow Solver 

for Unsteady Multiple Blade Row Turbomachinery Simulations,” ASME GT2001-0348, Jun, 2001.

41. To, W. M., “Unsteady Aero Computation of a 1-1/2 

Stage Large Scale Rotating Turbine”, NASA/CR—2012-

217438, 2012.

42. Dring, R.P., Fair, M.F., Joslyn, H.D., Power, G.D., and 

Verdon, J.M., “The Effect of Inlet Turbulence and Rotor/Stator Interactions on the Aerodynamics and Heat 

Transfer of a Large Scale Rotating Rig Turbine Model, 

Vol. I – Final Report,” NASA/CR—1987-4079, 1987.

43. Snyder, C.A., and Thurman, D.R.: “Effects of Gas Turbine Component Performance on Engine and Rotary 

Wing Vehicle Size and Performance,” NASA/TM—

2010-216907, Nov. 2010.

44. International Standard ISO 1940-1:2003/Cor. 1:2005(E), 

International Organization for Standardization, Switzerland, 2005, p. 11.

45. Skoch, G. J., Griffin, T. A., Stevens, M. A., Howard, S. 

A., Welch, G. E., and Giel, P. W., “Assessment of Variable-Speed Power-Turbine Component Test Capability 

at NASA Glenn Research Center,” NASA/TM—2012-

217422, 2012.
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-07-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Variable-Speed Power-Turbine Research at Glenn Research Center 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Welch, Gerard, E.; McVetta, Ashlie, B; Stevens, Mark, A.; Howard, Samuel, A.; Giel, Paul, 

W.; Ameri, Ali, A.; To, Waiming; Skoch, Gary, J.; Thurman, Douglas, R. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 877868.02.07.03.01.02.02 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18186 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2012-217605 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 02, 01, 03, 07, 34, and 09 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

The main rotors of the NASA Large Civil Tilt-Rotor (LCTR) notional vehicle operate over a wide speed-range, from 100 percent at takeoff 

to 54 percent at cruise. The variable-speed power turbine (VSPT) offers one approach by which to effect this speed variation. VSPT 

aerodynamics challenges include high work factors at cruise, wide (40° to 60°) incidence-angle variations in blade and vane rows over the 

speed range, and operation at low Reynolds numbers. Rotordynamics challenges include potential responsiveness to shaft modes within the 

50 percent VSPT speed-range. A research effort underway at NASA Glenn Research Center, intended to address these key aerodynamic and 

rotordynamic challenges, is described. Conceptual design and 3-D multistage RANS and URANS analyses, conducted internally and under 

contract, provide expected VSPT sizing, stage-count, performance and operability information, and maps for system studies. Initial steps 

toward experimental testing of incidence-tolerant blading in a transonic linear cascade are described, and progress toward 

development/improvement of a simulation capability for multistage turbines with low Reynolds number transitional flow is summarized. 

Preliminary rotordynamics analyses indicate that viable concept engines with 50 percent VSPT shaft-speed range. Assessments of potential 

paths toward VSPT component-level testing are summarized.

15. SUBJECT TERMS

Turbine aerodynamics; Rotordynamics; Turbine cascade; CFD; Gas turbine engine 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

30 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18