Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Douglas R. Thurman

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

Effects of Gas Turbine Component Performance

on Engine and Rotary Wing Vehicle Size and

Performance

NASA/TM—2010-216907

November 2010

U.S. ARMY

RESEARCH LABORATORY
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Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Douglas R. Thurman

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

Effects of Gas Turbine Component Performance

on Engine and Rotary Wing Vehicle Size and

Performance

NASA/TM—2010-216907

November 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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RESEARCH LABORATORY
Acknowledgments

We would like to thank the management of the Subsonic Rotary Wing vehicle project within the NASA Fundamental 

Aeronautics Program for supporting this research, and Mr. Michael Tong for his assistance using the

Weight Analysis of Turbine Engines (WATE) code for engine weight analyses.

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subject to revision as analysis proceeds.
NASA/TM—2010-216907 1

Effects of Gas Turbine Component Performance on Engine and 

Rotary Wing Vehicle Size and Performance 

Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Douglas R. Thurman 

U.S. Army Research Laboratory 

Glenn Research Center 

Cleveland, Ohio 

Abstract 

In support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project, further gas 

turbine engine studies have been performed to quantify the effects of advanced gas turbine technologies 

on engine weight and fuel efficiency and the subsequent effects on a civilian rotary wing vehicle size and 

mission fuel. The Large Civil Tiltrotor (LCTR) vehicle and mission and a previous gas turbine engine 

study will be discussed as a starting point for this effort. Methodology used to assess effects of different 

compressor and turbine component performance on engine size, weight and fuel efficiency will be 

presented. A process to relate engine performance to overall LCTR vehicle size and fuel use will also be 

given. Technology assumptions and levels of performance used in this analysis for the compressor and 

turbine components performances will be discussed. Optimum cycles (in terms of power specific fuel 

consumption) will be determined with subsequent engine weight analysis. The combination of engine 

weight and specific fuel consumption will be used to estimate their effect on the overall LCTR vehicle 

size and mission fuel usage. All results will be summarized to help suggest which component 

performance areas have the most effect on the overall mission. 

Nomenclature 

ft feet 

GW gross weight 

HP horsepower 

HPC high-pressure compressor 

HPT high-pressure turbine 

lbm pounds mass 

LCTR Large Civil Tiltrotor 

LCTR2 Large Civil Tiltrotor—iteration 2 

LPC low-pressure compressor 

LPT low-pressure turbine 

MRP maximum rated power 

NPSS Numerical Propulsion System Simulation 

OPR overall pressure ratio 

PR pressure ratio 

PSFC power specific fuel consumption, lbm/hour/HP 

sec second 

SOA State Of the Art 

T3 compression system exit temperature, °F 

T4 combustor exit temperature, °F 

°F degrees Fahrenheit 
NASA/TM—2010-216907 2

Figure 1.—Conceptual view of LCTR2. 

Introduction 

The NASA Heavy Lift Rotorcraft System Investigation (Ref. 1) identified a large tiltrotor 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 Tiltrotor—iteration 2) (Ref. 2) 

as seen in Figure 1. 

This vehicle iteration was designed to carry 90 passengers, at minimum cruising speed of 300 knots, 

1,000 nautical mile range. It is powered by four turboshaft engines designed for 7,500 HP each (MRP, 

sea level static conditions, standard day). Other design features included a rotor tip speed of 650 ft/sec in 

hover and 350 ft/sec during 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. 

Another contributor to efficient LCTR2 vehicle operations was an efficient and high power-to-weight 

gas turbine power plant. The previous study (Ref. 3) identified notional gas turbine characteristics and an 

engine layout that could meet fuel and weight requirements, while also identifying potential levels of 

component performance to meet overall vehicle objectives. However the inclusion of advanced 

technologies and better understanding component performance and designs for this particular engine size 

class might suggest different configurations that could better meet or exceed overall vehicle and mission 

requirements. This effort will therefore review the previous notional engine design, its component 

performance assumptions and revised baseline values for this effort. Then the analysis tools and 

methodology will be discussed, including some discussion on updated component performance 

assumptions and the inclusion of advanced compressor and turbine technologies. This will result in new 

engine configurations in terms of their specific size, weight and fuel consumption, which will then be 

used to estimate their effect on the overall vehicle for the baseline mission. 

Gas Turbine Engine Study, Previous and Updated Assumptions 

Previous engine efforts included several assumptions that helped define an optimum cycle (from a 

thermodynamic and power specific fuel consumption perspective). An example block representation of 

the turboshaft engine model is shown in Figure 2. Table 1 includes the sets of assumptions used in the 

previous parametric analyses and the updated baseline values. The previous assumptions came from a 

variety of sources (preliminary estimates from component technical experts and previous engine studies—

although not necessarily from the same engine class). Each particular component performance assumption 

will be discussed further in the next section. The engine as defined met all performance requirements, 

although several areas needing refinement were also noted. Overall parameters from the previous 

optimum engine and an operational engine (Rolls Royce AE1107, the modern turboshaft engine powering 

the Bell Boeing V-22 rotorcraft) are listed in Table 2 and which suggest conditions (higher temperatures 

and pressures, reduced corrected flow rates and total engine weight) that will require significantly 
NASA/TM—2010-216907 3

improved performance and possibly advanced technologies (aerodynamic and material/structural 

performance). Since this was a preliminary study, engine parameters were approximate values, to help 

define initial requirements for temperatures, pressures and applicable technologies at the relevant flow 

size. These would then be used to refine assumptions for the various component performances and the 

conceptual design (which was beyond the scope of the original work). Also during the previous effort, as 

the engine concept evolved from a single to two-spool gas generator core, the split of compression work 

performed by each spool of the compression system was assumed to be equal. The equal work split 

assumption will be revisited as part of this effort. 

Figure 2.—Block representation of a two-spool core turboshaft gas turbine model (from Ref. 3). 

TABLE 1.—GAS TURBINE ENGINE STUDY ASSUMPTIONS 

Engine parameter Prior-study assumptions (Ref. 3) Updated baseline 

(For this study) 

Compressor efficiency 85% polytropic efficiency 87% polytropic efficiency 

Compression work split 

(Horsepower-HPC/Horsepower-LPC) 1.0 (equal splits) 0.67 (optimized splits) 

Overall pressure ratio 10 to 60 10 to 60 

Combustor temperature (design/maximum), F 2000 to 3200 2400 to 3200 

Turbine efficiency 85% adiabatic efficiency 87% adiabatic efficiency 

Turbine cooling: metal temperature, °F +200 from internal, advanced, 

ultra high bypass engine 

Equal to internal, SOA high 

bypass engine 

Turbine cooling: Relative cooling flow factor 

(as defined in Ref. 9) 

Maintained from internal SOA 

turboshaft engine 

Maintained from internal SOA 

turboshaft engine 

TABLE 2.—GAS TURBINE ENGINE PARAMETERS 

Parameter AE1107 

(Used in original LCTR study) 

Notional engine 

(Concluded from LCTR2 study) 

Horsepower 6,000 7,500a

Weight, lbm 971 1,000a

Airflow, lbm/sec 35.3 30 

Power specific fuel consumption, lbm/hour/HP 0.426 0.37a

Overall pressure ratio (OPR) 16.7 30 

Compressor exit temperature, T3, °F 810 1099 

Combustor exit temperature, T4, °F 2200 3000 

Corrected flow: lbm/sec 

 Compressor entrance 

 Compressor exit 

35.5 

3.2 

30 

1.4 a

Parameters are from Reference 3. 
NASA/TM—2010-216907 4

Analysis Methodology 

The initial process was to develop improved performance estimates for the major engine components. 

These included compression efficiency, turbine efficiency and turbine cooling assumptions, which will be 

discussed later. With that information, a parametric assessment was performed varying engine overall 

pressure ratio (OPR) from 10 to 60, and combustor exit temperatures (T4) from 2400 to 3200 °F, to 

determine the optimum combination as defined as minimizing power specific fuel consumption (PSFC – 

lbm/hour/HP). Each engine was sized (determination of airflow) to produce 7,500 HP (the engine size 

from the LCTR2 study). The object-oriented analysis framework, the Numerical Propulsion System 

Simulation (Ref. 4) (NPSS), 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 diagram representative of a two-spool core (three-spool 

overall), turboshaft engine used within this effort is shown in Figure 3. As can be seen, compression is not 

only split between 2 spools (shafts), but the high pressure compressor (HPC) also is assumed to be an 

axial-centrifugal design. How the mix of compression between the axial and centrifugal components was 

determined will be discussed in the next section. Finally, utilizing the extensibility of NPSS, further 

elements were defined to drive specific parameters to desired values and insure continuity of mass, 

momentum and energy. 

The gas turbine flow path layouts and weights were generated using the Weight Analysis of Turbine 

Engines (Ref. 5) (WATE) program. Using the output from NPSS (mass flows, temperatures, pressures, 

velocities, etc.) and further user input, WATE sizes the various mechanical and flow components for the 

gas turbine engine, determining materials, dimensions and weights for the different components 

represented according to conceptual level design rules and stress analyses. 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 can form the basis for more detailed follow-on studies (which was done as reported in Refs. 6 

and 7). 

The previous NASA Heavy Lift Rotorcraft Systems Investigation (Ref. 1) also performed sensitivity 

studies which included estimates for the effect of changing engine weight and PSFC on the LCTR 

configuration gross weight and mission fuel. Those technology sensitivity results were used to provide a 

first-order estimate for the effects of different engine component technologies on vehicle and fuel weights 

and will be discussed later. 

Figure 3.—Block representation of NPSS turboshaft gas turbine model. 

 

Low Pressure 

Compressor

High Pressure 

Compressor -

Centrifugal

Combustor High Pressure 

Turbine

Low Pressure 

Turbine Inlet Nozzle

High Pressure Shaft Power Turbine Shaft

Bleeds for turbine cooling

Horsepower Extraction Drive system

T3 T4

Main Rotor

Power 

Turbine

Low Pressure Shaft

High Pressure 

Compressor -

Axial
NASA/TM—2010-216907 5

Results and Discussion 

In this section, assumptions for determining the baseline engine configuration while varying the 

compression system work split and centrifugal stage pressure ratio will be discussed, as will be the 

assumed baseline and advanced technology levels for the compression efficiency, turbine efficiency and 

turbine cooling. Then engine optimization results for each advanced technology applied separately and 

with all three technologies combined will be discussed. To conclude this section, the engine weights and 

PSFC results will be used to estimate effects on overall vehicle and mission sizing. 

Compression System Work Split 

In a single-spool engine, all compression is done with a single compressor that could consist of all 

axial stages or a mix of axial and a centrifugal stage. For such a device, the best mix of stage types comes 

from engine requirements (as determined by vehicle sizing and mission analysis) and compressor design 

effort (as is discussed in Ref. 6). For a two-spool gas generator core, the compression work is split 

between the high and low pressure spools; the choice of work split has some effect on the design and 

performance for the compression system and the overall engine. In the previous engine study, an equal 

work split was chosen between the two spools, assuming that the two stage turbine from the single-spool 

design would become one turbine stage for each spool. This assumption produced a viable, conceptuallevel solution. In this study, the work split was varied for the initial baseline engine to values of 

Horsepower-HPC spool/Horsepower-LPC spool of 1.5, 1.0, and 0.667. With the assumed HPC 

configuration of an a priori unknown number of axial stages and a final centrifugal stage, a further 

possible variable is the pressure ratio of the centrifugal stage. For an aft-centrifugal stage, entrance 

conditions are already very high in pressure and temperature, limiting the pressure ratio and efficiency 

potentially obtainable. Pressure ratios of 2.0 and 2.6 were chosen for this initial configuration analysis. 

An optimum cycle was determined for each assumed horsepower split; and its size and weight were 

estimated. An example result of the effect of T4 and OPR on PSFC is shown in Figure 4, assuming a 

compression work split of 0.667 and a centrifugal stage pressure ratio of 2.6. For that combination of 

compression work split and centrifugal pressure ratio, to obtain a valid HPC axial stage pressure ratio 

(PR > 1), engine OPR had to be greater than 25. After reviewing the optimum PSFC results, if the 

minimum appears to have occurred between 2 values of OPR used, the case was rerun at an OPR halfway 

and optimum OPR was chosen from among those 3 OPR values. For example, it often appeared that the 

optimum occurred between an OPR of 40 and 45. The case was rerun with an OPR of 42.5, and the 

optimum was chosen from the cases with an OPR of 40, 42.5, and 45 which resulted in the minimum 

PSFC value. 

Figure 4.—PSFC versus OPR and T4 at HPC centrifugal PR = 2.6, compression 

work split of 0.667. 

0.34

0.35

0.36

0.37

0.38

0.39

20 30 40 50 60

PSFC

OPR

T4=3200F

T4=3000F

T4=2800F

T4=2600F

T4=2400F
NASA/TM—2010-216907 6

The results for this analysis are shown in Table 3, with minimum values for PSFC and engine weight 

noted. Optimum cycles all converged to the maximum allowable T4 resulting in the same OPR, due to 

thermal efficiency benefits for the reference component efficiencies and cooling technology. To achieve 

the given horsepower of 7,500, engine airflow varied less than 4 percent (maximum to minimum), engine 

weight varied by just over 7 percent, and PSFC varied only by about 1 percent. For a given compression 

work split, minimum weight was always realized at a compression work split of 0.667. A centrifugal 

pressure ratio of 2.6 achieved minimum PSFC. Since the LCTR2 has a cruise-dominated mission, PSFC 

was considered to be the more critical figure of merit. This resulted in the chosen compression work split 

of 0.667 and a centrifugal pressure ratio of 2.6 for the rest of this study. This configuration had the best 

PSFC and the second best engine weight. Fixing these parameters also reduced the number of engines 

analyzed to a manageable effort. The WATE representation for this baseline cycle is shown in Figure 5. 

As engine parameters changed, flow path geometries did slightly change (as well as the number of axial 

stages in the HPC), but this figure is representative of the various engines with differing technology 

assumptions. The WATE program determined 7 axial stages for the LPC, with the HPC consisting of 

1 axial and 1 centrifugal stage. The number of axial stages was determined by maintaining an axial stage 

work factor that was consistent with the assumed level of efficiency for that component. The HPT and 

LPT each resulted in 1 stage, according to assumed study design rules, with a power turbine of 4 stages. 

TABLE 3.—ENGINE RESULTS FROM COMPRESSION WORK SPLIT AND CENTRIFUGAL PR 

Work split Centrifugal 

pressure ratio 

Airflow, 

lbm/sec 

OPR T4, 

°F 

PSFC, 

lbm/hr/HP 

Engine weight, 

lbm

1.0 2.0 28.6 42.5 3200 0.3470 971 

1.5 2.0 28.3 42.5 3200 0.3480 978 

0.667 2.0 29.0 42.5 3200 0.3465 913 

1.0 2.6 28.4 42.5 3200 0.3449 971 

1.5 2.6 28.1 42.5 3200 0.3464 943 

0.667 2.6 28.7 42.5 3200 0.3442 930 

Figure 5.—WATE flow path representation for the baseline engine. 

LPC HPC 

Power Turbine 

HPT LPT 
NASA/TM—2010-216907 7

Figure 6.—Compressor polytropic efficiency versus flow and technology level. 

Compression System Efficiency 

Compression system efficiency is an important parameter for any gas turbine cycle. Improved 

compressor efficiency has a two-fold effect on overall engine efficiency. First, it directly affects the 

amount of work that must be produced from the turbine sharing the same shaft; an improvement in 

compression efficiency reduces turbine work and pressure ratio requirements. This can improve the gas 

generator exhaust pressure that subsequently produces the work output from the engine. Second, an 

improvement in compressor efficiency also results in cooler compressor bleed flows that are used for 

turbine cooling, also reducing that parasitic loss to the engine and can result in an improvement in engine 

performance. These topics and other aspects about compressors with respect to the Rotary-Wing Project 

are discussed in Reference 8. In the previous effort, compressor polytropic efficiency was assumed to be a 

constant, regardless of flow size. Reference 8 looked at this in greater detail and resulted in notional 

compressor polytropic efficiency curves that were used for this study and are shown in Figure 6 (at exit 

corrected flows greater than 5 lbm/sec, the value for efficiency was maintained at that value). Thus the 

value for polytropic efficiency for the LPC, HPC axial stages and centrifugal stage were each set based on 

their respective exit corrected flows and assumed technology level. The effect of compression system 

efficiency on the optimum cycle parameters and overall engine PSFC and weight will be summarized and 

discussed at the end of this section. 

Turbine Efficiency 

Turbine efficiency is also an important parameter for any gas turbine cycle. Improved turbine 

efficiency enables further work potential for all components downstream in the gas turbine. Because 

turbines are downstream of the combustor, they are subjected to some of the highest temperatures found 

in the engine, while being required to maintain good performance and adequate life between maintenance 

periods. Since the previous engine study effort, there have been limited efforts to update turbine 

assumptions. Therefore, the levels of turbine efficiency were maintained from the earlier study with an 

assumption of 2 percent improvement (see Table 1), similar to the improvement in compression 

efficiency for future technology, to demonstrate the potential for improvement in engine performance and 

configuration with similar improvements in turbine performance. 

0.83

0.84

0.85

0.86

0.87

0.88

0.89

0.90

0.91

0.92

0123456

Polytropic efficiency

Exit corrected flow, lbm/sec

Future 

Technology

Present 

Technology
NASA/TM—2010-216907 8

Turbine Cooling 

Turbine cooling flows are a parasitic loss to the cycle and represent flow that is not within the main 

engine flow stream at all times, resulting in the loss of potential work from this flow. However, the use of 

turbine cooling flows enables higher combustion (and downstream) operating temperatures, resulting in 

increased engine efficiency and work output, while maintaining turbine life. The method of Gauntner 

(Ref. 9) is used to estimate the amount of cooling flow required, based on the main gas temperature and 

mass flow entering each turbine stage, the turbine cooling flow gas temperature, allowable turbine metal 

temperature and relative cooling flow factor (1.0 is the base, lower numbers indicate less cooling flow is 

required). The previous study assumed that allowable turbine metal temperatures for this engine class 

would achieve values 200 °F greater than those assumed for large engines, while maintaining cooling 

flow effectiveness. It has been suggested that this is too aggressive to develop within the assumed time 

period and the more difficult environment in this, smaller, engine class. A smaller engine effectively has a 

higher surface area for cooling per turbine main flow than larger engines; thus for a given technology 

level, smaller engines will have a higher, relative cooling flow factor (require more cooling flow as a 

fraction of the main flow). For this study, maximum turbine metal temperature was assumed to be 

1900 °F for rotors, 2000 °F for stators (same temperature levels assumed for present, large engines in 

NASA studies), while employing relative cooling flow factors previously used for small turboshaft engine 

internal studies (2.5 for rotors, 2.0 for stators). An additional, ultimate case was performed in which the 

cycle was reoptimized assuming no turbine cooling flow required. This could be illustrative of advanced 

materials and gives a potential of the performance benefit available if no turbine cooling were required. 

Optimum Engines Results From Inclusion of Various Advanced Technologies 

The results for the optimum engine are presented in Table 4. Significant improvements in PSFC and 

engine weight are possible employing the baseline assumptions on a new engine, with additional benefits 

realizable with advanced technologies. For each advanced technology assumption, optimum PSFC was 

realized at the highest T4 and some increase in engine OPR; the combination may be difficult to achieve 

for this engine size class. The conditions where each case optimized, combined with the small benefits 

from improved turbine cooling, suggest turbine cooling assumptions are probably too aggressive, and 

drive the cycle optimum to higher T4 and OPR than is truly potentially realizable. Thus, results from this 

effort highlight the need for more detailed turbine technology investigations before additional engine 

studies are warranted. 

TABLE 4.—OPTIMUM ENGINE SIZING RESULTS FROM INCLUSION OF VARIOUS ADVANCED TECHNOLOGIES 

Engine Airflow, 

lbm/sec 

OPR T4, 

°F 

PSFC Weight Delta 

PSFC, 

percent 

Delta 

weight, 

percent 

LCTR2 (Ref. 2) ----- ----- ----- 0.3700 1000.0 Base Base 

Baseline (updated assumptions) 28.7 42.5 3200 0.3442 948.5 –7.0 –5.1 

Baseline + Compressor efficiency 28.0 47.5 3200 0.3321 928.7 –10.2 –7.1 

Baseline + Turbine efficiency 28.7 45.0 3200 0.3364 942.9 –9.1 –5.7 

Baseline + Turbine cooling 27.1 45.0 3200 0.3407 933.9 –7.9 –6.5 

Baseline + Advanced turbine 

materials (“no turbine cooling”) 

19.5 60.0 3200 0.3144 902.1 –15.0 –9.8 

Baseline + All technologies 

(except “no turbine cooling”) 

26.4 52.5 3200 0.3223 918.7 –12.9 –8.1 
NASA/TM—2010-216907 9

Assuming material technologies that removed the requirement for turbine cooling altogether (no 

turbine cooling) realized the largest improvements in PSFC and engine weight. The significant reduction 

in gas generator airflow (–32 percent from the baseline) did reduce the gas generator size and weight, but 

the power turbine still had to derive a given amount of power from a significantly smaller and hotter gas 

flow. This resulted in the need for an additional power turbine stage for this engine configuration, versus 

the other technologies analyzed, which reduced the amount of weight benefit that was achieved. It should 

also be noted that for this configuration, at the high OPR, compressor exit conditions also impose 

stringent design constraints with exit conditions of 0.62 lbm/sec corrected flow and almost 1400 °F. The 

combination of all technology advancements realized almost the same benefits as the no turbine cooling 

case, showing that some improvements across several areas can realize significant gains. 

Effect of Engine Component Technology on Vehicle Gross and Mission Fuel Weights 

As part of the heavy lift rotorcraft investigation reported in Reference 1, LCTR vehicle sensitivities to 

different technology assumptions were generated and reported. Although the LCTR vehicle reported in 

Reference 1 is similar to the present vehicle iteration, there have been some changes in the design 

requirements and technology assumptions that resulted in the LCTR2 baseline design, which should also 

affect the magnitude of the LCTR2 sensitivities to changes in engine performance and weight. They were 

used here as an initial estimate, to help indicate potential areas for further analysis and refinement, as well 

as suggest areas to maximize technology investment. The effects of drive system weight and the 

combination of engine PSFC and power to weight assumptions on the LCTR vehicle are listed in Table 5. 

Between these two sets of sensitivities, one can estimate the effects of engine weight and PSFC 

separately, which are shown in Table 6 and discussed here. Since engine weights are only about 1/3 as 

heavy as the total drive system, changes in engine weights should only have 1/3 of the effect on the LCTR 

vehicle as changes in drive system weights. This effect can then be removed from the engine technology 

effects for changing both the engine power to weight ratio and PSFC and will result in a separate factor 

for only the change in PSFC. These new sensitivities were then used to estimate the change in gross 

weight and mission fuel for the inclusion of advanced engine technologies studied as part of this effort 

and are exhibited in Table 7. Improvements in PSFC enabled substantial reductions for fuel consumption 

and overall vehicle gross weight. Engine weights are already pretty low and a much smaller fraction of 

the total vehicle, so gains there were smaller, but there is still some effect on gross weight and fuel. Since 

the “no turbine cooling” approach enabled the largest reduction in PSFC and engine weight, it also leads 

in the reduction of vehicle gross weight and mission fuel. 

TABLE 5.—EFFECT OF ADVANCED ENGINE TECHNOLOGY ON LCTR (FROM REF. 1) 

Attribute Advanced 

technology 

(Baseline for 

LCTR) 

Current 

technology 

Percent change Overall effect of current 

technology on LCTR vehicle 

(Base is advanced 

technology) 

Drive system weight 

Weight technology 

factor 

0.67 1.00 +50% drive system weight 

(drive is 13% of gross 

weight) 

+23% Gross weight 

+21% Mission fuel 

Engine technology 

PSFC, lbm/hour/HP 0.3243 0.426 +31% fuel usage +23% Gross weight 

Power/Weight +70% Mission fuel 

(HP/lbm) 

7.48 6.49 +15% engine weight 
NASA/TM—2010-216907 10

TABLE 6.—ESTIMATED EFFECTS OF ENGINE WEIGHT AND PSFC ON LCTR VEHICLE 

(BASED ON RESULTS FROM REF. 1) 

Attribute Change Overall effect on 

LCTR vehicle 

Rationale 

Engine weight (only), 

from drive system 

weight 

+50% lbm/HP +8% GW 

+7% Mission fuel 

Engine weight is approximately 1/3 of drive 

system weight, ratio effect on LCTR from 

Table 5 

Combined engine 

weight and PSFC 

+31% PSFC 

+15% lbm/HP 

+23% GW 

+70% Mission fuel 

Same cumulative effect as Table 5 

Engine weight (only) +15% lbm/HP +2.4% GW 

+2.1% Mission fuel 

Effect of engine weight (only) using 

(15%/50%) lbm/HP ratio 

Engine PSFC (only) +31% PSFC +21% GW 

+68% Mission fuel 

Remove effect of engine weight from 

cumulative effect of both engine weight and 

PSFC 

TABLE 7.—ESTIMATED EFFECTS OF ADVANCED TECHNOLOGIES ON 

VEHICLE GROSS WEIGHT AND MISSION FUEL 

 Effect of PSFC improvements Effect of engine weight 

improvements 

Combined effect 

Engine Delta 

PSFC, 

percent 

Delta 

GW, 

percent 

Delta 

fuel, 

percent 

Delta 

eng.Wt, 

percent 

Delta 

GW, 

percent 

Delta 

fuel, 

percent 

Delta 

GW, 

percent 

Delta 

fuel, 

percent 

LCTR2 (Ref. 2)—base ---- ---- ---- ---- ---- ----- ----- ----- 

Baseline (updated assumptions) –7.0 –4.7 –15 –5.1 –0.8 –0.7 –5.5 –16 

Baseline + Compressor efficiency –10.2 –6.9 –22 –7.1 –1.1 –1.0 –8.0 –23 

Baseline + Turbine efficiency –9.1 –6.2 –20 –5.7 –0.9 –0.8 –7.1 –21 

Baseline + Turbine cooling –7.9 –5.4 –17 –6.5 –1.0 –0.9 –6.4 –18 

Baseline + Advanced turbine 

materials (“no turbine cooling”) 

–15.0 –10.2 –38 –9.8 –1.6 –1.4 –11.8 –39 

Conclusions 

Engine component performance estimates for a LCTR2-sized turboshaft engine were updated and 

new, optimum engine thermodynamic, flow path and weight analyses have been performed to help 

estimate the effects of engine component performance on engine size, performance and subsequently, 

vehicle size and mission performance. The inclusion of state of the art values from commercial turbofan 

engines and advanced small-engine technologies suggested that there are still significant specific fuel 

consumption gains to be realized. The “no turbine cooling” approach gave the largest benefit, but the 

combination of all technologies (compressor and turbine efficiency, and turbine cooling flow reductions) 

realized similar improvements and might be the more balanced and obtainable technology path. Engine 

weights are already a fairly small fraction of the vehicle gross weight, so gains there were small, but 

notable. Overall, this and the previous study (Ref. 3) consistently identify the optimum engine cycle as a 

two-spool gas generator core, operating below 30 lbm/sec mass flow, with elevated T4 (3000 to 3200 °F 

or more), and increased OPR, greater than 30 or even 40, for the LCTR2 engine class. While more 

detailed examination of the efficiencies and material capabilities is warranted, in concert with a full 

vehicle sizing analysis, these results describe alternative technology paths that can help refine future 

engine requirements for a LCTR. 
NASA/TM—2010-216907 11

References 

1. Johnson, W., Yamauchi, G.K., and Watts, M.E., “NASA Heavy Lift Rotorcraft Systems 

Investigation,” NASA/TP—2005-213467, September 2005. 

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. Snyder, C.A., and Thurman, D.R., “Gas turbine characteristics for a Large Civil Tilt-Rotor (LCTR),” 

AHS International, 65th Annual Forum & Technology Display, Grapevine, Texas, May 27–29, 2009, 

NASA/TM—2010-216089. 

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. Tong, M.T., and Naylor, B.A., “An Object-Oriented Computer Code for Aircraft Engine Weight 

Estimation,” GT2008–50062, ASME Turbo-Expo 2008, June 9–13, 2008, NASA/TM—2009-

215656. 

6. Veres, Joseph P., “Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements,” AHS 

International, 65th Annual Forum & Technology Display, Grapevine, Texas, May 27–29, 2009, 

NASA/TM—2009-215641. 

7. Veres, Joseph P., and Thurman, Douglas R., “Conceptual Design of a Two Spool Compressor for the 

NASA Large Civil Tilt Rotor Engine,” AHS International, 66th Annual Forum & Technology 

Display, Phoenix, Arizona, May 11–13, 2010. 

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

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

International, 65th Annual Forum & Technology Display, Grapevine, Texas, May 27–29, 2009. 

9. Gauntner, J.W., “Algorithm for Calculating Turbine Cooling Flow and the Resulting Decrease in 

Turbine Efficiency,” NASA TM 81453, 1980. 
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Effects of Gas Turbine Component Performance on Engine and Rotary Wing Vehicle Size 

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Snyder, Christopher, A.; Thurman, Douglas, R. 

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14. ABSTRACT

In support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project, further gas turbine engine studies have been performed 

to quantify the effects of advanced gas turbine technologies on engine weight and fuel efficiency and the subsequent effects on a civilian 

rotary wing vehicle size and mission fuel. The Large Civil Tiltrotor (LCTR) vehicle and mission and a previous gas turbine engine study 

will be discussed as a starting point for this effort. Methodology used to assess effects of different compressor and turbine component 

performance on engine size, weight and fuel efficiency will be presented. A process to relate engine performance to overall LCTR vehicle 

size and fuel use will also be given. Technology assumptions and levels of performance used in this analysis for the compressor and turbine 

components performances will be discussed. Optimum cycles (in terms of power specific fuel consumption) will be determined with

subsequent engine weight analysis. The combination of engine weight and specific fuel consumption will be used to estimate their effect on 

the overall LCTR vehicle size and mission fuel usage. All results will be summarized to help suggest which component performance areas 

have the most effect on the overall mission. 

15. SUBJECT TERMS

Turboshaft; Gas turbine engine; Performance prediction; Weight 

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