NASA/TM-2010-216089

Gas Turbine Characteristics for a

Large Civil Tilt-Rotor (LCTR)

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Douglas R. Thin-man

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

February 2010
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NASA/TM-2010-216089

Gas Turbine Characteristics for a

Large Civil Tilt-Rotor (LCTR)

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Douglas R. Thin-man

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

Prepared for the

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

sponsored by the American Helicopter Society

Grapevine, Texas, May 27-29, 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

February 2010
This report is a formal draft or working

paper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary findings,

subject to revision as analysis proceeds.

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.

Level of Review: This material has been technically reviewed by technical management

Available from

NASA Center for Aerospace Information National Technical Information Service

7115 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161

Available electronically at http://gltrs.grc.nasa.gov
Gas Turbine Characteristics for a Large Civil Tilt-Rotor (LCTR)

Christopher A. Snyder

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Douglas R. Thunman

U.S. Army Research Laboratory

Glenn Research Center

Cleveland, Ohio 44135

Abstract

In support of the Fundamental Aeronautics Program,

Subsonic Rotary Wing Project, an engine system study has

been undertaken to help define and understand some of the

major gas turbine engine parameters required to meet

performance and weight requirements as defined by earlier

vehicle system studies. These previous vehicle studies will be

reviewed to help define gas turbine performance goals.

Assumptions and analysis methods used will be described.

Performance and wei ght estimates for a few conceptual gas

turbine engines meeting these requirements will be given and

discussed. Estimated performance for these conceptual

engines over a wide speed variation (down to 50 percent

power turbine rpm at hi gh torque) will be presented. Finally,

areas needing further effort will be suggested and discussed.

Nomenclature

C.G. center of gravity, inches

eff efficiency

fps feet per second

ft feet

hp horsepower

HPC high-pressure compressor

HPT high-pressure turbine

lbm pounds mass

LCTR Large Civil Tilt Rotor

LP low pressure

LPC low-pressure compressor

LPT low-pressure turbine

Max maximum

N actual speed, rpm

Nc corrected speed, NIJO , rpm

PR pressure ratio

PSFC power specific fuel consumption, lbm/hr/lip

PT power turbine

sec second

T3 compression system exit temperature, °F

T4 combustor exit temperature, °F

Vtip rotor tip velocity, feet per second

W actual mass flow, lbm/sec

We corrected mass flow, W * FO/8 , lbm/sec

Wturb power turbine actual mass flow, lbm/sec

ratio of actual to standard pressure

ratio of actual to standard temperature

Introduction

The NASA Heavy Lift Rotorcraft System Investigation

(Ref. 1) identified a large tilt rotor as the best concept to meet

the various airspace and other requirements for the future,

short-haul regional market. This evolved into a conceptual

vehicle designated as LCTR2 (Large Civil Tilt Rotor—

iteration 2) (Ref. 2) as seen in Figure 1

This vehicle iteration was designed to carry 90 passengers

at 300 knots with at least a 1,000 n nu range; powered by four

turboshaft engines designed for 7,500 hp each. Other design

features included a rotor tip speed of 650 ft/sec in hover and

350 ft/sec durin g cruise, enabled by a two-speed gearbox. This

range of rotor tip speeds was needed to achieve the high level

of performance and efficiency at two very different flight

conditions. The rotor tip speed variation could theoretically be

obtained using a variable diameter rotor or multiple-speed

gearboxes (orVa combination of these or other approaches);

this work is focusing on achieving all speed variation from the

engine. Although the exact requirements and characteristics

for such a vehicle class are still being researched, performing

analyses on a representative vehicle will help understand the

sensitivities for such a design, help guide research efforts to

reduce risks, and develop a suite of technologies from which

this new vehicle class and capability can evolve and be

developed. The final vehicle design could use one or a

combination of these variable rotor tip speed concepts,

determined from the vehicle's specific design and nussion

requirements and the state of these various required

technologies.

NASA/TM-2010-216089
Figure 1.—Conceptual view of LCTR2.

This report details gas turbine engine technology

assumptions and analyses performed to estimate the engine

parameters needed to obtain sufficient performance to meet

operational goals for the proposed vehicle concept. This

vehicle would use the turboshaft version of a gas turbine

engine. The core gas turbine engine develops high energy

(high temperature and pressure) gas to power a separate power

turbine and shaft that supplies horsepower and torque to the

main drive system. Initial cycle parametrics were performed to

suggest gas turbine engine characteristics that would meet

LCTR2 performance requirements. Based on these overall

characteristics, therniodynamic engine analyses assumin g a

one-spool core (all compression on one shaft) and a two-spool

core (with the compression split between two sets of

compressors, each on their own shaft) were performed. As

stated above, both configurations have a separate power

turbine and shaft. For the thermodynamic analyses,

compressor component performance maps were generated and

used to estimate off-design performance at the hover and

cruise flight points. At the cniise condition, gas turbine engine

performance was estimated at 100, 75 and 50 percent power

turbine rpm to quantify engine performance at reduced power

turbine speeds. Results will be discussed and suggestions for

farther analysis will be given. Follow-on studies are underway

to perform more detailed analyses of compression and turbine

systems and component performance. It is expected that the

more detailed component analyses will be incorporated and

reported in subsequent studies and reports. It must be noted

that this work is not expected to identify the specific gas

turbine engine attributes and cycle definitions to meet all

requirements. Its purpose is to further refine requirements and

identify possible components, systems or subsystems that

could enable such new classes of engine and vehicle designs

and operations while also uncovering areas requiring further

exploration and development.

Analysis Methodology

Engine system studies were performed to estimate the

major gas turbine engine parameters that would meet

performance requirements for the previously defined vehicle

and mission. Engine power, weight and fuel consumption are

important performance parameters that will help define

possible engine configurations. As part of the parametric

analysis, compressor pressure ratio (PR) was varied from 5 to

60, assuming a constant polytropic efficiency of 88 percent.

Although turbomachinery efficiency would vary depending on

engine size and configuration, that effect was deferred to later

analyses. Combustor exit temperature was varied from 2000 to

3200 °F (in 400° increments). To get turbine cooling bleed

estimates for the core turbines, the method of Gauntner

(Ref. 3) was used, assuming metal temperatures of 2200 °F

for the stator, 2100 °F for the rotor. These turbine metal

temperatures are higher than present, small gas turbines to

include the effects of incorporation of improved turbine

material temperature capabilities and cooling techniques in

these smaller turbine sizes. The power turbine was assumed to

be urncooled. These should be reasonable temperatures and

efficiencies for engines in the LCTR2-size class with entry-inservice in roughly the 2020 timeframe.

The object-oriented analysis framework, the Numerical

Propulsion System Simulator (NPSS) (Ref. 4), was used to

perform the gas turbine analyses. NPSS contains standard 0/1-

D elements for the gas turbine components. These are

configured into a representative steady-state, thermodynamic

model. An example block dia gram representative of a onespool core, turboshaft is shown in Figure 2. Further elements

are defined to drive specific parameters to desired values and

insure continuity of mass, momentum and ener gy. After initial

cycle parameters were determined, CMGEN (Ref. 5) was used

to generate compressor performance maps that would be more

representative of the flog-speed characteristics for a given

compressor PR and size during the off-design analyses.

The gas turbine flow path and weight were generated using

the WATE (Ref. 6) program. Using the output from NPSS

(mass flows, temperatures, pressures, velocities, etc.) and

further user input, A"ATE sizes the various mechanical and

flow components for the gas turbine engine, determining

materials, dimensions and weights for the different

components represented. As part of the process, WATE also

produces a graphical representation that can be used to check

for reasonable component dimensions and ensure that there

are no discontinuities or sharp turns in the gas flow path. The

results of these analyses also form the basis for more detailed

follow-on studies.

Results and Discussion

Initial Engine Parametrics

The preliminary analysis is instructive to suggest engine

parameters to meet vehicle and mission requirements. Figure 3

shows calculated Power Specific Fuel Consumption (PSFClbm/hr fuel per hp produced) versus high pressure compressor

(HPC) PR and combustor exit temperature (T4). Also included

in the graph are areas representative of the performance

regions for the Honeywell Aerospace T5.5 and GE Aviation

NASA/TM-2010-216089
HPC bleeds for turbine cooli

Highpressure HighPower Inlet Duct Duct com- Diffuser Combustor pressure Duct Nozzle

pressor I. turbine turbine

High-pressure shaft

Low-pressure shaft

T3 T4 Horsepower Drive ♦ Main

extraction system rotor

Figure 2.—Block representation of a one-spool core turboshaft gas turbine model.

0.600

CL 0.550

E

0.500

0

0.450

0

0.400

U_

0.350

0 10 20 30 40 50 60

HPC pressure ratio

Figure 3.—Power Specific Fuel Consumption (PSFC, Ibm/hr/hp) versus compressor

pressure ratio and combustor exit temperature, T4.

T700 families of turboshaft engines, the version of Rolls

Royce AE 1107 in the Bell Boeing V-22 aircraft (all from

Ref 7) and the estimated level of PSFC used in the LCTR2

system studies. With the assumed technology levels, there is a

minimum in the PSFC curves at a HPC PR of 30 (for values of

T4 around 2800 to 3200 °F—both curves fall almost on top of

each other). To meet the LCTR2 design of 7,500 hp per engine

in that region of T4 and HPC PR would require an airflow of

approximately 30 lbm/sec. Therefore, based on the technology

assumptions, the engine parameters (at the sea level static

design) used for subsequent analyses were: airflow of

30 lbrri/sec, overall pressure ratio of 30, and T4 of 3000 °F.

Table 1 compares the major parameters for this gas turbine to

the Rolls Royce AE1107 (the engine in the Bell Boeing V-22

rotorcraft, a modern turboshaft engine in a slightly smaller

power class). To meet the PSFC requirements; the notional

study engine will need to operate at temperatures and

pressures significantly higher than those found in present

turboshaft engines; but at levels already in modern, large

engines; while maintaining compressor and turbine efficiency

and performance at the required much smaller airflows and

blade sizes. A few conceptual engine configurations were

investigated and will now be discussed in concert with their

representative engine thermodynamic and flow path

arrangements.

TABLE 1.—GAS TURBINE ENGINE PARAMETERS

Parameter AE1107 Notional engine

Horsepower 6,000 7,500,

Weight, Ibm 971 1,000,

Airflow, lbm/sec 353 30

PSFC, lbun/hr/hp .426 .3 7'

Overall pressure ratio 16.7 30

Compressor exit temperature, T3, °F 810 1099

Combustor exit temperature, T4, T 2200 3000

Corrected flow:

Compressor entrance 35.5 30

Compressor exit 3.2 1.4

'Parameters are from Reference 2.

One-Spool Core Engine

The first engine to be modeled was a one-spool core (with a

free turbine on a second spool) with an all-axial compressor.

This allowed the engine flow path to be laid out and to

generate compressor component perfonnance maps to do an

initial check on off-design performance. Compressor maps

were generated at the design pressure ratio and flow.

Two nussion profiles were evaluated, each with a hover

condition and a range of conditions at cruise. Gas turbine

performance was calculated at the key points (hover and initial

cruise power), to verify sufficient engine horsepower

was available at full power turbine rpm (hover) and at

NASA/TM-2010-216089
100 percent, 7.5 percent and the reduced power turbine

mechanical rpm identified for optimum rotor performance at

cruise. Results are given in Table 2 detailing the flight

condition and pertinent turbomachinery operating

characteristics. This preliminary analysis indicated that the gas

turbine was able to achieve horsepower requirements for all

flight points except one. There was insufficient engine

horsepower for the mission 2 initial cruise point (operating at

higher than the 300 knot requirement at reduced rotor / power

turbine rpm). However, the engine was able to meet vehicle

power requirements after about 1/3 of cruise fuel was

consumed, reducing the vehicle weight and power needed. It

was also found that at the cruise condition, 100 percent power

turbine mechanical speed actually results in an aerodynamic

over speed as a result of the drop in ambient temperature with

altitude. This suggests that even if a multispeed gearbox was

used, power turbine rpm could need to be reduced for cruise

operation.

TABLE 2.-OPERATION OF ONE-SPOOL CORE TURBOSHAFT ENGINE OVER

REFERENCE MISSION PROFILES AT VARIOUS POWER TURBINE RPMS

Altitude, Speed. Vtip, fps AT. Meet Core rpm HPC HPT Ne T4, °F PT Ne PIP/ We-PT

ft knots (rotor) °F ]]p`' (actual,/design) Nc Wtrrb° (lbnl/s)

0 0 650 +27 Yes 1.02 99.2 1 100.4 3091 98.5 1 261.7 12.54

Mission 1

2.000 0 650 -45 Yes 0.99 95.6 101.0 2864 101.8 215.9 12.16

(100% PT rpm) 28,000 303.4 350 0 Yes 0.90 98.0 99.9 2367 111.7' 210.2 12.60

(75% PT rpm) Yes 0.90 98.1 100.3 2352 84.1'' 208.4 12.86

(50% PT rpm) Yes 0.91 99.3 100.5 2405 59.9 198.9 13.21

Mission 2

5,000 0 650 -36 Yes 0.99 97.0 100.6 2870 101.9 227.1 12.31

(100% PT rprn) 28,000 330 350 0 Yes 0.92 99.1 99.7 2464 109.7" 228.7 12.71

(75% PT rpni) Yes 0.92 99.4 100.2 2459 82.5 225.5 13.00

(50% PT spin) NO 0.93 100.0 100.4 3480 59.1 308.8 13.26

(50% PT rpm) End 28,000 330 350 0 Yes 0.92 99.9 100.4 2472 59.2 207.9 13.26

'Actual power turbine horsepower /mass flow (hp/lbm/sec)

bMultispeed gearbox required to match rotor and power turbine speeds)

The operating points for the core turbomachinery and

power turbine are shown in Figures 4 to 6. As can be seen in

Figure 4 and Figure 5, the core compressor and turbine operate

in a fairly narrow band over the fli ght profile. As shown in

Figure 6, the power turbine has a much larger variation in its

rpm. Although power turbine corrected flow varies less than 6

percent from its sea level design value ; its pressure ratio

increases about 1.5 percent from the full speed rpm at hover to

the part speed rpm of the cruise. As shown in Table 2, the

amount of engine power required per pound of actual power

turbine mass flow is fairly constant for both missions at both

flight points. This was not unexpected, since flight altitude and

speed are determined by the engine power available (among

other factors), with actual engine airflow and power falling

with the increase in altitude and the accompanying drop in

ambient density_ However; it was hoped that power

requirements per lbm airflow would drop to allow power

turbine rpm to decrease as the mission progressed from hover

to cruise, such that the main rotor rpm could also be reduced

to maintain high rotor efficiency. For this analysis, at reduced

rotor and power turbine rpm, the loss of turbine efficiency is

compensated by an increase in pressure ratio to maintain

power production. Maintaining a constant power turbine

horsepower output per lbm airflow with an almost 50 percent

reduction in rpm without loss of efficiency requires turbine

design unique to this vehicle and nussion class. A preliminary

analysis looking at variable vanes and blades to meet this

requirement is given in Reference 8. One such design concept

includes power turbine variable incident nozzles, incident

tolerant blades and additional stages. Whatever the design

solution, further analysis is needed for the power turbine to

verify its performance at almost constant power to weight flow

over such a range of rotational speeds while still meeting

engine weight goals.

Engine gas flow path dimensions and weights were

generated assuming an all axial compression system, output

from the WATE analysis is shown in Figure 7. It resulted in an

engine plus accessories total weight of 946 lb (about 5 percent

lighter than assumed in the LCTR2 study). At this engine

airflow and horsepower class with one spool, all axial

compression results in 10 stages to achieve the desired

compression (tip speeds are below 1,000 ft per second for the

last 4 stages). This specific application also results in very

small blade heights for the latter stages (less than 1 in. for the

last 4 stages, the last stage blade height is only 0.57 in.). This

size blade could be a challenge for achieving and maintaining

efficiency over the life of the engine. Further discussion on

this particular design challenge is given in Reference 9. The

power turbine WATE results assume typical turbine stage

design. If additional power turbine stages using incident

tolerant blades (and other variability) were needed to maintain

good operation over its large speed range, these factors could

increase power turbine weight.

NASA/TM-2010-216089 4
T^ 0.84

1.0

1.4 0.78

1.8

2.2

2.6

110 0.72

0.66

0.60

0.54

0.48

0.42

int

0.36

40

3f

3(

0

2,1

m

m

a 2(

d

1(

5 10 15 20 25 30 35

Corrected flow

Figure 4.-One-spool core turboshaft engine compressor performance map and operating points.

0.9

0.8 _^-

•

0.7

0.6 .

-- - --- - - - --

w 0.5

1

-------

0.3 Design point n

0.2 - - - - 1 2 3 4 5 6 7

Pressure ratio

2.4 -

2.2

i - --

3: 10

-0 1.8 (1 = 65 ::. 30

m 50

a 1.6 ^y Map scalars, - 70

o

! We = 0.0296 : 90 U 1.4 Eff = 0.940 - 110

1.2 n Design point PR = 0.324

N = 0.0175 • 130

150

1.0 1 2 3 4 5 6 7

Pressure ratio

Figure 5.-One-spool core turboshaft engine turbine performance maps and operating points

NASA/TM-2010-216089 5
U.0

0.8

0.7

7 0.6

0.5

0.4

w 0.3

0.2

0.1

00

1 2 3 4 5 6 7 8

Pressure ratio

14

12

10

11 :'" R = 74 - 30

10 Map scalars,

9- 0 We = 0.070

C) Eff = 0.933 '90

8 ; 110 PR = 0.252 130 7 - ° • Design point N = 0.0259 150

6

1 2 3 4 5 6 7 8

Pressure ratio

Figure 6.-One-spool core turboshaft engine power turbine performance maps and operating points.

9

9

16

12

8

4

0

-4

-8

-12

-16

Weights Dimensions

Bare engine weight 895.7 Engine length 63.1

Accessories weight 50.0 Engine pod C. G. 21.9

Engine weight 945.7 Engine max diameter 24.0

Inlet and nacelle weight 96.4 Nacelle max diameter 28.7

Total engine pod weight 1042.1 Total engine pod length 69.3

-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7

Figure 7.-One-spool core turboshaft engine WATE output.

NASA/TM-2010-216089 6
Two-Spool Core Engine

Noting the challenging compressor design for a one-spool

core, a two-spool core (with a free power turbine on a

third spool) was modeled. A block representation is given in

Figure 8. Compression was split with roughly equal enthalpy

change for the low and high pressure compressors, which

resulted in a low pressure compressor (LPC) pressure ratio of

9.2 and a high pressure compressor (HPC) pressure ratio of

almost 3.3 (maintaining the same engine overall pressure ratio

of 30). Results are given in Table 3 detailing the flight

condition and pertinent operating characteristics for the low

pressure spool turbomachinery and power turbine. This

preliminary analysis indicated that the two-spool core

turboshaft engine could achieve the horsepower requirements

for all flight points (and power turbine mechanical rpms).

Operation of the high pressure spool is essentially at constant

corrected rpm and correct mass flow rates (less than 2.3

percent variation in the compressor and less than 1 percent in

the turbine aerodynamic values). The high pressure spool had

a very minor physical over-speed (1 percent above the design

value) at hover, and was only operating at 90 to 97 percent of

the physical design speed for cruise (the difference in

corrected or aerodynamic versus physical rpm is from the

reduction in component entrance temperatures going from hotday hover to standard-day altitude conditions).

The operating points for the low pressure spool

turbomachinery and power turbine are shown in Figures 9 to

11. As seen in Figure 9, the low pressure compressor corrected

speed varies a maximum of 7 percent from the design

condition; this maximum point is a result of the engine

matching response to the hover condition on a hotter than

average day (45 °F hotter at 2;000 ft altitude). As shown in

Figure 10, there is also essentially no variation in the low

pressure turbine corrected mass flow (<1 percent), althou gh its

pressure ratio varies slightly (-4 to 6 percent from the design

value). The operation of the power turbine is very similar

between this two-spool core engine (Figure 11) and the onespool core version (Figure 6). This was expected with the

similarity in vehicle power requirements and gas properties

(mass flow, temperature and pressure) provided by the core of

each engine, therefore, the previous analysis still applies.

TABLE 3.-OPERATION OF TWO-SPOOL CORE TURBOSHAFT ENGINE OVER

REFERENCE MISSION PROFILES AT VARIOUS POWER TURBINE RPMS

Altitude, Speed, Vtip, fps AT, Meet LP shaft rpm LPC LPT T4, PT HP,/ We-PT

ft knots (rotor) F lip? (achral/design) Nc Nc °F Nc Wturb^ (Ibnl/s)

0 0 650 -27 Yes 1.02 99.2 100.1 3103 98.3 263.9 11.96

Mission 1

2.000 0 650 -45 Yes 0.97 93.6 99.5 2826 102.3 208.5 11.51

(100% PT 1prn) 28,000 303.4 350 0 Yes 0.89 96.7 99.9 2319 112.7' 201.8 11.95

(75% PT rpm) Yes 0.89 96.9 100.6 2295 85.1' 200.2 12.18

(50% PT rpm) Yes 0.91 98.4 101.4 2333 60.8 190.5 12.54

Mission-')

5.000 0 650 -36 Yes 0.97 95.3 99.7 2817 102.7 218.5 11.66

(100% PT rprn) 28,000 330 350 0 Yes 0.91 98.0 100.1 2413 110.8' 219.5 12.06

(75% PT rpm) Yes 0.91 98.5 100.8 2397 83.5' 216.3 12.32

(50% PT rpm) Yes 0.93 100.3 101.6 2455 59.4 205.6 12.62

(50% PT rn) End 28,000 330 350 0 Yes 0.92 99.2 101.5 2397 60.1 199.1 12.60

'Actual power turbine horsepower hnass flow (HP/lbm/sec)

'Multispeed gearbox required to match rotor and power turbine rpnrs.

HPC bleeds for turbine cooling

Figure 8.-Block representation of a two-spool core turboshaft gas turbine model.

NASA/TM-2010-216089
12-

10 -I

0 8

m

V

a 6

4^

2 -I

0.84

1.0

1.4

1.8 0.78

2.2

2.6

110 0.72

05

0.66

0.60

0.54

0.48

In point

^- 0.42

5 10 15 20 25 30 35

Corrected flow

Figure 9.-Two-spool core turboshaft engine low-pressure compressor performance map and

operating points.

U. t)

0.8

0.7

0.6

U

0.5

0.4

w 0.3

0.2

0.1

00

1.0 1.5 2.0 2.5 3.0 3.5 Pressure ratio

5.5

0 4.5 10

4.0 R = 74 30 0 3.5- y Map scalars, _ _ 70

o Wc = 0.0267 90 v 3.0 / Eff = 0.923 a 110

2.5 n Design point PR = 0.972 0 130 N = 0.0194 150

2.0

1.0 1.5 2.0 2.5 3.0 3.5 Pressure ratio

Figure 10.-Two-spool core turboshaft engine low-pressure turbine performance maps and operating points

NASA/TM-2010-216089 8
n Design point

2 3 4 5 6 7 8 9 10

Pressure ratio

/i

10

R = 74 ....... 30

50 Map scalars, 70 We = 0.067

Eff = 0.933 ~ 90' 110 PR =0.24 m 130 N = 0.026 150

n Design point

0.9

0.8

0.7

0.6

0.5

0.4

w 0.3

0.2

0.1

0.0

1

13

12

11

10

0

W 9

U

8

0

U 7

6

5

....

a' --------

i

1 2 3 4 5 6 7 8 9 10

Pressure ratio

Figure 11.—Two-spool core turboshaft engine power turbine performance maps and operating points.

Engine gas flow path dimensions and weights were

generated assuming an all axial compression system; output

from the WATE analysis is shown in Figure 12. It resulted in

an engine plus accessories total weight of 891 lb (about

11 percent lighter than assumed in the LCTR2 study).

Separating compression into two separate spools allowed

some reduction in compressor and turbine weight for the core

engine, although the core total shaft weight increased (the sum

of the two, lighter core shafts required weighed 80 lb more

than the one core shaft of the one-spool concept). It also

enabled a slightly larger compression exit blade height

(0.65 in., almost 14 percent higher). Very small, latter stage

blade heights are still an issue for this concept. This two-spool

core concept also adds the complexity of a third shaft, which

might be an issue considering the limitations on the third shaft

diameter size available imposed by the other two shafts and

torque. The third shaft might not be an issue for an aft turbine

power assembly, if such an option can be reasonably

integrated with the drive system.

To get away from the small axial stages at compression

system exit, another engine model was investigated in which

the axial hi gh pressure compressor from the previous WATE

iteration was replaced with a centrifugal stage. Althou gh this

centrifugal stage is operating at higher temperatures than

centrifugal stages in present engines; it is worthwhile to

understand centrifugal stages with increased temperature

capability mi ght enable good solutions to the LCTR2 engine

requirements- As shown in Figure 13, it is very similar to the

two-spool core, all axial, except for the hi gh pressure

compressor and combustor arrangement. Its weight at 1051 lb

is a bit more than the other two engine concept weights,

exceeding the LCTR2 wei ght requirement by 5 percent, and

alleviates the small axial compressor blade height issue. The

additional weight versus the two-spool, all axial engine is

from the centrifugal compressor and combustor. Conservative,

heavy materials were assumed, based on the high temperature

environment for these components. This additional weight

could be mitigated through better materials or careful engine

flow path design, but those efforts are deferred to follow-on

studies. The engine layout also has the same potential issue

with its third shaft as previously discussed.

NASA/TM-2010-216089 9
-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7 73.3

14

12

10

8

6

4

2

0

-2

-4

-6

-8

-10

-12

-14

Weights Dimensions

Bare engine weight 841.1 Engine length 64.6

Accessories weight 50.0 Engine pod C. G. 17.8

Engine weight 891.1 Engine max diameter 24.0

Inlet and nacelle weight 96.8 Nacelle max diameter 28.8

Total engine pod weight 987.9 Total engine pod length 70.8

14

12

10

8

6

4

2

0

-2

-4

-6

-8

-10

-12

-14

-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7 73.3

Figure 12.-AII axial, two-spool core turboshaft engine WATE output.

Weights Dimensions

Bare engine weight 1000.5 Engine length 62.5

Accessories weight 50.0 Engine pod C. G. 19.1

Engine weight 1050.5 Engine max diameter 24.0

Inlet and nacelle weight 97.6 Nacelle max diameter 28.8

Total engine pod weight 1148.1 Total engine pod length 68.8

Figure 13.-Axicentrifugal high-pressure compressor, two-spool core turboshaft engine WATE output.

NASA/TM-2010-216089 10
Conclusions

An engine thermodynamic, flow path and weight analysis

has been performed to help define relevant areas of interest to

achieve engine performance to meet LCTR2 performance

goals and enable a new vehicle class and operations into its

particular niche in the national airspace. Studies suggest that

modern; large gas turbine engine temperatures and efficiencies

are needed for the next generation of 7,500 hp class turboshaft

engines. There has already been some recent work to look at

different compression configurations for this specific engine.

Further efforts are needed to guide research to develop and

verify designs for axial and centrifugal compressors to operate

at temperatures, efficiencies, and sizes beyond the present

state of the art. Turbine technology will also be needed to

maintain high efficiencies, reduced cooling requirements, and

especially verify efficient operation over a wide power turbine

speed range.

References

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

Lift Rotorcraft Systems Investigation," NASA/TP-2005-

213467, Sep. 200.5.

2. Acree, C.W., Hyeonsoo, Y., and Sinsay, J.D., "Performance

Optimization of the NASA Large Civil Tiltrotor," International

Powered Lift Conference, London, UK. July 22-24, 2008.

3. Gauntner, J.W., "Algoritlnn for Calculating Turbine Cooling

Flow and the Resulting Decrease in Turbine Efficiency," NASA

TM-81453, 1980.

4. Jones, Scott M., "An Introduction to Thermodynamic

Performance Analysis of Aircraft Gas Turbine Engine Cycles

Using the Numerical Propulsion System Simulation Code,"

NASA/TM-2007-214690.

5. Converse, G.L.; and Giffin, R.G.: Extended Parametric

Representation of Compressor, Fans and Turbines Voltune I -

CMGEN User's Manual. NASA Contractor Report 174645,

March 1984.

6. Tong, M.T., Naylor, B.A., "An Object-Oriented Computer Code

for Aircraft Engine Weight Estimation," GT2008-50062, ASME

Turbo-Expo 2008, June 9-13, 2008.

7. Gunston, Bill, ed.: Jane's Aero-Engines. Issue 14, Jane's

Information Group Limited, Coulsdon, Surrey, 2003.

8. Chen, Shu-cheng, "Preliminary Axial Flow Turbine Design and

Off-design Performance Analysis Methods for the Rotary Wing

Aircraft Engines; II-Applications," AHS International,

65th Annual Forum & Technology Display, Grapevine, TX,

May 27-29, 2009.

9. Veres, Joseph P., "Compressor Study to Meet Large Civil Tilt Rotor

Engine Regturements," AHS International, 65th Annual Fonnn &

Technology Display, Grapevine, TX, May 27-29, 2009.

NASA/TM-2010-216089 11
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Gas Turbine Characteristics for a Large Civil Tilt-Rotor (LCTR)

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14. ABSTRACT

In support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project ; an engine system study has been undertaken to help

define and understand some of the major gas turbine engine parameters required to meet performance and wei ght requirements as defined by

earlier vehicle system studies. These previous vehicle studies will be reviewed to help define gas turbine performance goals. Assumptions

and analysis methods used will be described. Performance and weight estimates for a few conceptual gas turbine engines meeting these

requirements will be given and discussed. Estimated performance for these conceptual engines over a wide speed variation (down to 50

percent power turbine r m at high for ue) will be presented. Finally, areas needin g ftuther effort will be suggested and discussed.

15. SUBJECT TERMS

Turboshafts; Gas turbine engine, Performance prediction, Weight

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