Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Defi ning Gas Turbine Engine Performance

Requirements for the Large Civil TiltRotor (LCTR2)

NASA/TM—2013-218101

December 2013
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Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Defi ning Gas Turbine Engine Performance

Requirements for the Large Civil TiltRotor (LCTR2)

NASA/TM—2013-218101

December 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

67th Annual Forum and Technology Display (FORUM 67)

sponsored by the American Helicopter Society International (AHS)

Virginia Beach, Virginia, May 3–5, 2011
Acknowledgments

The author would like to thank the management of the Subsonic Rotary Wing vehicle project within the NASA Aeronautics 

Research Mission Directorate, Fundamental Aeronautics Program for supporting this effort. The author would also like to thank 

Cecil Acree, Wayne Johnson, Chris Silva, Jeff Sinsay, and Doug Thurman for the baseline LCTR2 vehicle model and NDARC help.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

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

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 fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2013-218101 1 

Defining Gas Turbine Engine Performance Requirements for 

the Large Civil TiltRotor (LCTR2) 

Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Defining specific engine requirements is a critical part of 

identifying technologies and operational models for potential 

future rotary wing vehicles. NASA’s Fundamental 

Aeronautics Program, Subsonic Rotary Wing Project has 

identified the Large Civil TiltRotor (LCTR) as the 

configuration to best meet technology goals. This notional 

vehicle concept has evolved with more clearly defined mission 

and operational requirements to the LCTR-iteration 2 

(LCTR2). This paper reports on efforts to further review and 

refine the LCTR2 analyses to ascertain specific engine 

requirements and propulsion sizing criteria. The baseline 

mission and other design or operational requirements are 

reviewed. Analysis tools are described to help understand their 

interactions and underlying assumptions. Various design and 

operational conditions are presented and explained for their 

contribution to defining operational and engine requirements. 

These identified engine requirements are discussed to suggest 

which are most critical to the engine sizing and operation. The 

most-critical engine requirements are compared to in-house 

NASA engine simulations to try to ascertain which operational 

requirements define engine requirements versus points within 

the available engine operational capability. Finally, results are 

summarized with suggestions for future efforts to improve 

analysis capabilities, and better define and refine mission and 

operational requirements. 

Nomenclature 

CRP Contingency rated power 

HOGE Hover, out of ground effect 

IRP Intermediate rated power 

ISA International standard atmosphere 

LCTR2 Large Civil TiltRotor—iteration 2 

MCP Maximum continuous power 

MRP Maximum rated power 

NDARC NASA Design and Analysis of Rotorcraft 

NPSS Numerical Propulsion System Simulation 

OEI One engine inoperative 

OGE Out of ground effect 

PR Pressure ratio 

PT Power turbine 

SLS Sea level static 

T3 Compression system exit temperature, °F 

T4 Combustor exit temperature, °F 

Vbr Velocity at best range 

VSPT Variable-speed power turbine 

W Mass flow, lbm/sec 

Wcorr Corrected mass flow, W   , 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 tiltrotor as the best concept to meet 

the various airspace and other requirements for the projected, 

future, short-haul regional market. This evolved into a 

conceptual vehicle designated as Large Civil TiltRotor—

iteration 2 (LCTR2) (Ref. 2) as seen in Figure 1.

This iteration of the conceptual vehicle was designed to 

carry 90 passengers, at a minimum cruising speed of 300 knots 

and 1,000 nautical mile range. It is powered by four turboshaft 

engines designed for 7,500 HP each [maximum rated power 

(MRP), at sea level static (SLS), International standard 

atmosphere (ISA)]. 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. Subsequent 

studies (Refs. 3 to 5) have suggested that minimum vehicle 

gross weight (indicative of development and purchase costs) 

and minimum mission fuel (related to operational costs) might 

be achieved at slightly higher rotor cruise tip speeds, with 

Reference 3 suggesting a weak minimum around 400 ft/sec. 

While these studies have further and more clearly defined the 

vehicle size and overall vehicle operation, results have not 

been reported from an engine-centric point of view. Reporting 

specific engine requirements clarifies the engine performance 

assumptions in the vehicle analyses, while also enabling the 

engine technologists and designers to better understand and 

match requirements. For this study, the LCTR2 vehicle and 

mission model from Reference 3, assuming the 400 ft/sec 

rotor cruise tip speed, was further exercised to elicit further 
NASA/TM—2013-218101 2 

Figure 1.—Conceptual view of LCTR2. 

engine performance and vehicle sensitivity to changes in 

engine performance assumptions. 

This report gives details for the baseline vehicle and 

mission, focusing on engine assumptions and requirements. 

The analysis tools are briefly described, then results from the 

vehicle and mission sizing and analysis code from an engine 

requirements point of view are presented. The engine 

performance points are compared to data from a NASA inhouse engine performance code, further refining engine 

performance requirements. Results are summarized and future 

plans are discussed. 

Vehicle/Mission Requirements 

The LCTR2 mission requirements/summary are given in 

Table 1 (from Ref. 3) and except for takeoff and landing 

operations, are similar to a regional jet mission. Table 2 lists a 

matrix of takeoff conditions used to check engine power levels 

and examine their effect on vehicle sizing. Takeoff at the 

high/hot condition had the highest engine power requirement; 

while takeoff from sea level, standard day had a slightly higher 

mission fuel load. Both takeoff conditions were set as 

constraints for the vehicle sizing. Landing at 5,000 ft, ISA+ 

20 °C was chosen for all missions, which has the highest 

landing engine power requirement, although it was still 

significantly less than takeoff. Another operational requirement 

is one engine inoperative (OEI); the most critical condition 

would be during low speed departure just after takeoff, when 

the rotorcraft is at its heaviest without sufficient forward speed 

to augment lift from the wings. During this OEI event, the 

vehicle must execute a vertical landing as soon as possible with 

just three of its four engines operational. 

The assumed engine power rating structure is an additional 

vehicle/mission sizing parameter. The engine power available 

over the various vehicle and mission sizing points is directly 

linked to the engine design and an engine’s power rating 

structure. Typical engine ratings are used within this study and 

their definitions are listed in Table 3. At engine power levels 

above maximum continuous power (MCP), the engine is 

operating at more stringent temperatures and pressures, which 

limit engine useful life. The greater the increase in power, the 

more stress on engine components and reduced time that such 

power would be available (before engine maintenance or 

replacement would be required). The balanced use of these 

higher engine power ratings can result in a smaller vehicle, 

fuel load and engine size to meet mission requirements, while 

still meeting reasonable maintainability or life. Typical aircraft 

missions include engine power levels up to maximum rated 

power (MRP) for takeoff, although this power level is only 

maintained within the airport boundaries and reduced to MCP 

or less for community noise and engine life considerations. 

The rest of the mission is typically at engine power levels of 

MCP or less, except for landing. For the 1 min hover, out of 

ground effect (HOGE) and landing, engine power levels up to 

MRP may be used if conditions warrant (such as when the 

vehicle is still near maximum weight, and/or high, hot 

conditions). 

There are additional considerations in defining and sizing a 

vehicle concept, including relevant missions and other 

constraints. The baseline vehicle used for this study, as well as 

additional sizing and design constraints, are discussed further 

Reference 3. The optimum gross weight design and simulation 

models from that study were used as the starting point for this 

effort; baseline LCTR2 vehicle characteristics are given in 

Table 4. 

TABLE 1.—LCTR2 MISSION REQUIREMENTS 

Mission summary Approximate 

time in mode 

Takeoff + 2 min hover OGE 5,000 ft ISA+20 °C 2 min 

Climb at Vbr (credit distance to cruise segment)a 20 min 

Cruise at Vbr for at least 1000 nm range, 

28,000 ft, ISA 

3 hr 

Descend at Vbr (no range credit) 20 min 

1 min hover OGE + landing, 5,000 ft, 

ISA+20 °C 

1 min 

Reserve (diversion): 100 nm Vbr, 28,000 ft, ISA 22 min 

Reserve (emergency): 30 min Vbr, 5,000 ft, 

ISA+20 °C 

30 min 

a

Climb segment includes possible engine-rotor gear ratio shift event. 

TABLE 2.—TAKEOFF ALTITUDE/ 

TEMPERATURE DESIGN MATRIX 

Altitude Temperature 

5,000 ft ISA + 20 °C 

2,000 ft ISA + 25 °C 

0 ft (SLS) ISA + 25 °C 

0 ft (SLS) ISA + 0 °C 

TABLE 3.—TYPICAL ENGINE POWER RATINGS 

Rating Power rating Time limit Power/MCP 

MCP Max continuous none 1.000 

IRP Intermediate rated 30 min 1.185 

MRP Maximum rated 10 min 1.269 

CRP Contingency rated 2.5 min 1.330 
NASA/TM—2013-218101 3 

TABLE 4.—BASELINE LCTR2 CHARACTERISTICS 

Design Value 

Cruise speed (90 percent MCP), knots 300 

Rotor radius, ft 31.4 

Engine power (MCP, SLS, ISA), hp 46,476 

Gross weight, lb 96,500 

Payload (90 pax), lb 19,800 

Structure, lb 29,000 

Fuselage group—9,900 lb --------- 

Rotor group—7,400 lb --------- 

Wing and empennage—7,600 lb --------- 

Gear, engine nacelle, other—4,100 lb --------- 

Systems and Equipment, lb 14,600 

Engines and drive systems, lb 14,100 

Mission fuel, lb 13,900 

Other, lb 5,100 

Analysis Methodology 

The NASA Design and Analysis of Rotorcraft (NDARC) 

code was used to perform the vehicle design and sizing with 

multiple requirements and constraints. Reference 6 documents 

the background behind NDARC and its complete theory. Its 

theoretical basis and architecture is given in Reference 7. The 

program includes models for the major components and 

systems of rotorcraft vehicles (such as fuselage, rotors, wings, 

tails, and propulsion). These components can be further 

refined into subcomponents, defining performance and weight 

attributes that the program combines into a complete model 

for a given aircraft. During this process, various mission, 

sizing or operational requirements can be defined that will size 

(or resize) various aircraft components. For a defined (or 

sized) aircraft, performance can also be calculated for alternate 

missions or conditions. As part of this effort, alternative 

mission and propulsion attributes were assigned to the 

previous defined LCTR2 vehicle to further understand engine 

requirements for various aspects of its design and operational 

mission and the effect of changing propulsion capabilities on 

its sizing. 

Within NDARC, the default method for turboshaft engine 

performance “is based on curve-fits of engine performance 

data for existing or projected engines over a range of operating 

conditions” (Ref. 6). The engine model uses referred 

conditions to correct power, fuel consumption and net jet 

thrust for altitude, non-standard temperatures and operation at 

power turbine (PT) rpm at other than design conditions. 

Engine performance with technology assumptions equivalent 

to technology levels for the rest of the LCTR2 vehicle were 

used. Mission modeling assumes the use of multispeed 

gearbox systems to maintain engines at or close to optimum 

PT rpm. 

Transition From Hover to Cruise Rotor Speed 

A new requirement on the engine and drive system is being 

formulated to address possible constraints during rotor 

transition from hover to cruise rpm. Once the vehicle converts 

from helicopter to airplane mode and gains some flight speed, 

the wings are able to augment lift, significantly reducing rotor 

power requirements. Reducing rotor rpm from the hover to the 

cruise condition could further reduce community noise, but 

could increase drive system torque requirements (and weight) 

if rotor power levels are not reduced from the takeoff 

condition. If rotor cruise tip speed reduction is achieved by 

using multi-speed gearbox/transmissions (which enable the 

engines to operate most efficiently around a single speed), 

there will be a transition period during which the gear ratio 

between the engine PT and rotor is changed. An example 

block diagram of the drive system with the rotors, gearboxes, 

engines and speed changing modules is shown in Figure 2 

(from Ref. 4). 

Figure 2.—Drive system block diagram with speed changing modules. 

191 rpm 191 rpm

3.1 : 1

3 : 1

2.5 : 1

3.1 : 1

3 : 1

2.5 : 1

1.54 :1 8,000 rpm 1.54 :1 

5,194 rpm 5,194 rpm

MID-WING 

GEARBOX FOR 

ACCY POWER

TILT AXIS GEARBOX

GEN

PUMP

2.4 : 1

or

4.4 : 1 

15,000 rpm 15,000 rpm

2.4 : 1

or

4.4 : 1 

15,000 rpm 15,000 rpm

PROP ROTOR GEARBOX

3.5:1 3.5:1

Same engine configuration 

as shown on the left, but 

out-of-plane 

Same engine configuration 

as shown on the right, but 

out-of-plane 

Speed 

Changing 

Module

191 or 103 rpm 191 or 103 rpm
NASA/TM—2013-218101 4 

Figure 3.—Block representation of three spool turboshaft engine. 

An initial procedure for the speed shift process is proposed 

here: Slow the rotors from hover to cruise tip speed (maintain 

rotor power levels to not exceed drive system torque levels—

probably defined by takeoff, HOGE). Each engine is operating 

at less than optimum PT rpm (which can limit engine power 

and even cause engine instability). For this design, there is 

insufficient power for only one engine to power each rotor (if 

the vehicle is to maintain a reasonable climb profile or during 

cruise). Therefore, it is further assumed to use the crossshafting between the two rotors to maintain constant rotor 

power during the shift process. There is a flight envelope 

(altitude, speed and climb state) where one engine at a time 

could be reduced in power (the remaining three engines could 

potentially increase power to maintain constant rotor power 

levels), to enable that engine to shift from the hover to cruise 

gear ratio. Maintaining constant rotor power levels would 

enable the vehicle to continue its flight state, making the 

LCTR2 integration more seamless in the airspace. This initial 

effort will focus on describing power requirements for the 

mission starting at 5,000 ft, ISA + 20 °C, shifting at 6,000 ft 

(1,000 ft above the takeoff height) with rotor tip speed 

transitioning from 650 to 400 ft/s, while the LCTR2 is in 

airplane mode and early in the climb profile (where engine 

power availability should be the highest). Other cases were 

run, but this case requires the highest engine power levels 

(because of altitude and hot day). Investigation of shifting 

later in the mission profile (i.e., later or near the end of climb, 

at the start of cruise, etc.) will be deferred to later efforts. 

Variable-speed power turbine (Ref. 8) (VSPT) technology is 

also being explored to facilitate or possibly eliminate the 

gearbox/transmission shift. VSPT technology can be used to 

increase the PT operating region of highest power availability 

while maintaining efficiency. There is an increasing PT weight 

penalty as the PT high power availability operating region 

increases and minimizing efficiency losses for operation 

further from the PT optimum design speed. Some system-level 

analyses that have been performed assuming VSPT 

technology for the LCTR2 and their effect on the vehicle 

design are discussed in References 4 and 5. 

Engine Modeling 

A 1-D, steady-state, turboshaft engine model using 

technology levels equivalent to the LCTR2 was developed 

using the Numerical Propulsion System Simulation (NPSS) 

(Ref. 9) code and reported in References 10 and 11. A block 

diagram for this engine model is shown in Figure 3. To meet 

the stringent fuel efficiency requirements, the turboshaft 

engine would be a high overall pressure ratio design (overall 

pressure ratio around 40), with expected maximum combustor 

exit temperature at or over 3000 °F. To achieve this high 

pressure ratio, the core engine is envisioned with two separate 

shafts (or spools), in addition to a “free” PT on its own shaft, 

connected to the drive system to drive the rotor. Component 

maps were generated for the compression and turbine systems, 

which are used within the model to estimate off-design 

performance. This model was used to generate engine 

performance at selected flight conditions from the NDARC 

mission and sizing analyses to perform a more detailed 

understanding of engine operation and performance margins. 

Results and Discussion 

Engine Power Levels and Sizing 

CRP was the default, engine power rating level used for the 

OEI sizing condition, assumed at roughly 5 percent more 

power available than at MRP, and 33 percent more than MCP. 

For the variety of mission and constraints assumed for the 

LCTR2, OEI at the high/hot conditions (5,000 ft, ISA + 20 °C) 

was the most stringent engine power constraint and therefore 

determined engine size. It makes sense that the high, hot 

condition was the most stringent, engine power lapses more 

with increased altitude (5,000 ft, ISA + 20 °C) than for a 

hotter, but lower altitude (2,000 ft, ISA + 25 °C). A vehicle 

sized for OEI at 5,000 ft, ISA + 20 °C could meet OEI at 

2,000 ft, ISA + 25 °C with the three remaining engines 

operating at MRP levels and at just above the IRP level for 

SLS, ISA + 25 °C. To determine the sensitivity of vehicle size 

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—2013-218101 5 

TABLE 5.—EFFECT OF ENGINE POWER ON LCTR2 SIZING 

OEI power level Engine size, 

hp, SLS MRP 

Mission 

fuel, 

lb 

Gross 

weight, 

lb 

Max. possible takeoff weight,

lb 

133 percent MCP (base) 6,480 13,900 96,550 137,500 

127 percent MCP 7,000 14,500 99,320 146,400 

119 percent MCP 7,900 15,500 104,020 161,100 

TABLE 6.—MISSION/FLIGHT POINTS FOR ENGINE POWER 

Condition Speed, 

knots 

Altitude,

ft 

Temperature Total HP used/available Engine rating 

OEI 0 5,000 ISA + 20 °C a17,150/15,450 CRP 

HOGE/takeoff 0 5,000 ISA + 20 °C 16,550/19,610 MRP 

Climb 1 (shift) 172 6,000 ISA + 20 °C b11,084/11,240 75 percent MCP 

Climb 2 (MCP) 207 8,000 ISA + 20 °C 10,340/10,350 MCP 

Cruise start 284 28,000 ISA 7,620/11,070 MCP 

Cruise end 272 28,000 ISA 6,710/10,940 MCP 

Reserve-diversion 270 28,000 ISA 6,610/10,920 MCP 

Reserve-emergency 218 5,000 ISA 6,610/19,640 MCP 

HOGE/land 0 5,000 ISA + 20 °C 13,890/19,610 MRP 

a

OEI modeling assumes 111 percent CRP availability. 

b

Power available is MCP for only three engines (assumes one engine is shifting). 

for different levels of engine power capability beyond MCP 

was performed. The importance of higher power availability 

beyond MCP (even for limited times, such as 2.5 min for the 

OEI requirement) on vehicle sizing can be seen in Table 5. 

Meeting the OEI constraints with reduced levels of 

contingency power availability (beyond MCP) significantly 

increases engine sizing, mission fuel and vehicle gross weight. 

With all those penalties, there are some benefits to these 

vehicles increased engine sizes (and power). Such vehicles 

would then have greater maximum possible takeoff weight 

capability (calculated at MRP engine power, for SLS at ISA) 

and have even greater climb capability (but climb performance 

met or exceeded minimum climb requirements for all cases). 

Engine/Flight Requirements 

To define the flight state and engine power requirements, 

this paper focuses on what the author considers the most 

relevant points during a typical mission (summarized in 

Table 1). These flight points and power requirements are in 

Table 6 and come from the vehicle sizing and mission 

analyses performed by NDARC. The OEI requirement at 

5,000 ft, ISA + 20 °C (at the default CRP of 133 percent 

MCP) is a critical sizing point for the engines and the vehicle. 

For the OEI calculation, the methodology from Reference 2 

was used, paraphrased here: assuming a 10 percent reduction 

for OEI hover power, modeled as 111 percent (100/ 

90 percent) increase in power available as a practical 

approximation. HOGE was assumed to determine the takeoff 

condition (HOGE is more stringent as ground effects reduce 

power required). The next flight point could be the transition 

from hover to airplane mode (assumed to occur shortly after 

leaving the takeoff area). But with any forward speed, the 

accompanying lift would reduce engine requirements from the 

HOGE/takeoff requirements, so that flight point was omitted. 

Climb was modeled in two segments; the first segment is 

during rotor tip speed transition from hover to cruise and is 

assumed to occur 1,000 above the takeoff point (while still 

assuming hot day conditions). With any gearbox/transmission 

shift occurring here, only 75 percent of engine MCP is 

available (three engines out of four). Engine conditions were 

also generated before and after shift; before the shift, the PT 

would be at non-optimum conditions (just to note any 

differences in engine conditions between these two operating 

conditions). The second climb segment is after transition, 

continuing ascent to top-of-climb (all four engines would be 

capable of MCP, at optimum PT rpm). Power at the start and 

end of cruise defines the range of power required for the 

constant altitude cruise segment. The two reserve segments 

(for diversion and emergency) are included, but should not 

require high power levels. Power requirements for rotor speed 

transition from cruise back to hover rpm will not be discussed, 

but would probably occur during the low-power descent and 

would also not require high power levels (especially since the 

bulk of the mission fuel was consumed during the cruise 

portion and the vehicle is significantly lighter). The final point 

is HOGE and landing at the high/hot condition (5,000 ft, ISA 

+ 20 °C), but since the vehicle is again significantly lighter 

without the bulk of the mission fuel, this point should be 

within engine MRP capabilities. 

Engine Requirements: Relating Mission Results 

(NDARC) to Engine Performance Deck (NPSS) 

Engine data was generated for these flight points using the 

previously mentioned NPSS turboshaft model. This data was 

used to verify engine performance (power) and internal 

parameters (mass flows, turbomachinery operating conditions, 
NASA/TM—2013-218101 6 

pressures and temperatures) that help define the various 

performance requirements. Even with the wide range of flight 

points, plotting the turbomachinery operating points on their 

specific compressor and turbine performance maps suggests 

that the turbomachinery operates in a fairly small portion of 

their operating envelopes to meet LCTR2 power requirements, 

as can be seen in Figure 4 to Figure 9. In fact, the highpressure compressors (Figure 5 and Figure 6), high-pressure 

turbine (Figure 7) and low-pressure turbine (Figure 8) are 

operating at close to or essentially at constant aerodynamic 

conditions. The PT (Figure 9) does operate over a greater 

range of its operating envelope than the other components, but 

much of the greatest variance occurs at low power levels. The 

most notable is the “Reserve Emergency” condition, a low 

altitude diversion to another airport. For such conditions, 

engine capability is much greater than required; engines are 

operating at significantly less than MCP. The PT is also 

operating significantly off-design, but able to meet the low 

power requirements. 

Figure 4.—Low-pressure compressor performance map. 

Figure 5.—High-pressure axial compressor performance map. 
NASA/TM—2013-218101 7 

Figure 6.—High-pressure centrifugal compressor performance map. 

Figure 7.—High-pressure turbine performance map. 
NASA/TM—2013-218101 8 

Figure 8.—Low-pressure turbine performance map. 

Figure 9.—Power turbine performance map. 

 
NASA/TM—2013-218101 9 

Results from the engine performance analyses, at typical 

engine points at SLS and for the operating conditions listed in 

Table 6 are shown in Table 7. The OEI requirement at high, 

hot conditions determines engine size (for the assumed level 

of CRP), and that requirement is so demanding, that the 

engines are able to meet all other mission power requirements 

at or under their respective engine rating power levels (either 

MRP or MCP). After takeoff, engines stay at MCP combustor 

exit temperature levels through the climb profile (much like 

regional aircraft); the compressors experience some 

temperature reductions, mainly from altitude effects. There 

was sufficient engine power available for the transition from 

hover to cruise rotor tip speeds; before the shift, the engine PT 

is at non-optimum conditions, which slightly increases engine 

power setting (to get the same power output), internal 

conditions and specific fuel consumption. Also during that 

climb segment, right after transitioning to the lower rotor tip 

speed, drive system torque levels could be slightly higher than 

the HOGE takeoff torque levels if the engines are at MCP, 

which suggests one would have to be careful how quickly (and 

how much) to slow the rotors (or set engine power levels) as 

altitude increases (and engine power falls) to stay within drive 

system torque limits. Cruise requirements are fairly 

reasonable; temperatures are significantly less than takeoff or 

climb levels. There is actually sufficient cruise power 

available to run on only three engines and still be below 

engine MCP limits (cruise power levels are from 62 to 

67 percent of MCP—with four engines operating). The first 

reserve leg is at the cruising altitude, so it looks very similar to 

conditions at end of cruise. The second reserve leg was a latein-flight diversion at low altitude (5,000 ft, ISA, at velocity for 

best range). At such a low altitude, there is a lot of available 

power and the engine meets power requirements at very low 

power levels. At such low power levels, the turbomachinery is 

operating away from design conditions, which results in 

reduced component efficiencies and higher power specific fuel 

consumption. For the HOGE landing requirement, since the 

vehicle is significantly lighter than takeoff without most of the 

mission fuel, only 85 percent of the HOGE takeoff power 

level is required, with significantly lower combustor 

temperature (approximately 140 °F cooler), operating 

effectively below MCP conditions (although it could be 

running up to MRP levels). One last thing to note from the 

SLS engine power rating conditions is the effect of 

diminishing returns for the increase in available power with 

increase in combustor temperature (and amount of time such 

power levels would be available). 

TABLE 7.—ENGINE PERFORMANCE PARAMETERS 

Condition Airflow, 

lb/s 

Compression exit 

temperature, 

°F 

Compression 

exit pressure, 

psia 

Combustor exit 

temperature, 

°F 

Delivered 

power, 

HP 

Power specific fuel 

consumption, 

lb/hour/HP 

Standard engine rating points, SLS 

MCP 22.6 1136 546 2980 5105 0.359 

IRP 24.2 1185 600 3135 6050 0.349 

MRP 24.9 1205 623 3200 6476 0.345 

CRP 25.3 1226 639 3267 6787 0.345 

Flight points from Table 6 

OEI 19.8 1237 500 3248 5156 0.350 

HOGE/takeoff 18.1 1163 443 3046 4141 0.363 

Climb 1 (shift)a 17.6 1157 425 2983 3691 0.382 

Climb 1 (shift)b 17.2 1142 412 2937 3695 0.365 

Climb 2 (MCP) 14.4 1068 332 2712 2583 0.390 

Cruise start 9.4 911 209 2500 1902 0.324 

Cruise end 8.9 881 195 2411 1675 0.333 

Reserve-diversion 8.8 878 193 2400 1649 0.334 

Reserve-emergency 13.8 922 296 2311 1656 0.471 

HOGE/landing 16.9 1128 404 2905 3474 0.377 

a

Before shift, PT not at optimum rpm (61.5 percent of hover mechanical rpm) 

b

After shift, PT at optimum rpm (100 percent mechanical rpm) 
NASA/TM—2013-218101 10 

Conclusions 

Additional mission analyses were performed on the LCTR2 

vehicle from an engine-centric point of view; in hopes of 

further understanding and refining engine power requirements 

and their effect on the LCTR2 vehicle sizing and mission. 

During this effort, a process was described for the transition of 

the rotors from hover to cruise rpm. Based on the assumed 

mission, engine ratings and constraints, the OEI requirement 

at 5,000 ft, ISA+20 °C was the most stringent condition that 

defined engine power requirements and had the greatest effect 

on the vehicle and engine sizing. Lower levels of CRP beyond 

MRP and MCP would increase vehicle weight, mission fuel 

and engine sizing, but would potentially enable higher climb 

rates or a greater maximum takeoff weight. Results from the 

NPSS turboshaft engine simulation suggests that under the 

assumed engine power ratings, constraints, and missions that 

there are very limited conditions where engines are required to 

operate beyond MCP levels, with most conditions comfortably 

at or below this level. Engine turbomachinery components 

operate within a fairly narrow portion of their overall 

operational envelopes, except for the PT; but even its 

operating range was well within its performance envelope. 

After transition from helicopter to airplane mode, there is 

sufficient power available to maintain constant rotor power 

and climb rate while shifting the gear ratio between the 

engines and rotors to better match engine PT rpm for the 

slower, rotor cruise tip speed. For some conditions, depending 

on the amount of rotor tip speeds reduction and assumed 

power used during climb (MCP), torque levels under these 

conditions would be the higher than the typical torque limit 

design, increasing drive system torque requirements and 

weight (minor operational constraints could alleviate such 

occurrences). For the assumed turbomachinery performance 

and efficiencies, diminishing increases in power beyond MCP 

are achieved for further increases in combustor temperature, 

with significant reductions in engine life. These must be 

balanced with vehicle sizing and performance, and potentially 

other items that are beyond the scope of this study. 

Future Efforts 

Optimum rotor cruise tip speed is highly dependent on 

many interconnected assumptions. Recent NASA in-house and 

contracted efforts have tried to further refine engine and drive 

system performance and weights; this new understanding 

needs to be folded into this work. VSPT technology can 

change the shape and range of engine performance; new drive 

system designs and material technologies could change the 

weight and efficiency trends for a given speed reduction from 

engine PT to vehicle rotor. Combined, these may lead to a 

different vehicle and drive system design and operation to 

minimize vehicle gross weight and mission fuel and need to be 

applied to the LCTR2 design and optimization. These efforts 

are on-going and are within present research plans. Other 

missions and requirements continue to be defined that may 

lead to configurations other than the tiltrotor. With generic, 

new technologies and learning being applied, capabilities to 

model these vehicles faster and with greater fidelity offer the 

opportunity to find and define the next generation of rotary 

wing vehicles to help people get there faster, easier, and 

cheaper. 

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, Also, NASA/TM—2008-215359. 

3. Acree, C.W. Jr., “Integration of Rotor Aerodynamic 

Optimization With the Conceptual Design of a Large 

Civil Tiltrotor,” AHS Aeromechanics Conference, San 

Francisco, CA, January 20–22, 2010. 

4. Snyder, Christopher, A.; Robuck, Mark; Wilkerson, 

Joseph; Nordstrom, Carl., “Summary of the Large Civil 

Tiltrotor (LCTR2) Engine Gearbox Study,” NASA/TM—

2010-216908, International Powered Lift Conference, 

Philadelphia, Pennsylvania, October 5–7, 2010. 

5. Robuck, Mark; Wilkerson, Joseph; Zhang, Yiyi; Snyder, 

Christopher, A.; Nordstrom, Carl., “Design Study of 

Propulsion and Drive Systems for the Large Civil 

Tiltrotor (LCTR2) Rotorcraft,” American Helicopter 

Society 67th Annual Forum, Virginia Beach, VA, May 3–

5, 2011. 

6. Johnson, W., “NDARC, NASA Design and Analysis of 

Rotorcraft,” NASA/TP—2009-215402, December 2009. 

7. Johnson, W., “NDARC—NASA Design and Analysis of 

Rotorcraft: Theoretical Basis and Architecture,” AHS 

Aeromechanics Specialists’ Conference, San Francisco, 

California, January 2010. 

8. Ford, A., Bloxham, M., Turner, E. Clemens, E., and 

Gegg, S., “RTAPS VSPT Contract NNC10BA14B, 

Design Optimization of Incidence-Tolerant Blading 

Relevant to Large Civil Tilt-Rotor Power Turbine 

Applications,” NASA/CR—2010-217016, Nov., 2010. 

9. 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. 

10. Snyder, C.A., and Thurman, D.R., “Effects of gas turbine 

component performance on engine and rotary wing 

vehicle size and performance,” AHS International, 66th 

Annual Forum & Technology Display, Phoenix, AZ, May 

11–13, 2010. 

11. 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. 

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Defining Gas Turbine Engine Performance Requirements for the Large Civil TiltRotor 

(LCTR2) 

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Snyder, Christopher, A. 

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14. ABSTRACT

Defining specific engine requirements is a critical part of identifying technologies and operational models for potential future rotary wing 

vehicles. NASA’s Fundamental Aeronautics Program, Subsonic Rotary Wing Project has identified the Large Civil TiltRotor (LCTR) as the 

configuration to best meet technology goals. This notional vehicle concept has evolved with more clearly defined mission and operational 

requirements to the LCTR-iteration 2 (LCTR2). This paper reports on efforts to further review and refine the LCTR2 analyses to ascertain 

specific engine requirements and propulsion sizing criteria. The baseline mission and other design or operational requirements are reviewed. 

Analysis tools are described to help understand their interactions and underlying assumptions. Various design and operational conditions are 

presented and explained for their contribution to defining operational and engine requirements. These identified engine requirements are 

discussed to suggest which are most critical to the engine sizing and operation. The most-critical engine requirements are compared to inhouse NASA engine simulations to try to ascertain which operational requirements define engine requirements versus points within the 

available engine operational capability. Finally, results are summarized with suggestions for future efforts to improve analysis capabilities, 

and better define and refine mission and operational requirements.

15. SUBJECT TERMS

Turboshaft; Tiltrotor; Design; Gas turbine engine; Performance prediction 

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