Eric S. Hendricks and Michael T. Tong

Glenn Research Center, Cleveland, Ohio

Performance and Weight Estimates for an 

Advanced Open Rotor Engine

NASA/TM—2012-217710

September 2012

AIAA–2012–3911
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Eric S. Hendricks and Michael T. Tong

Glenn Research Center, Cleveland, Ohio

Performance and Weight Estimates for an 

Advanced Open Rotor Engine

NASA/TM—2012-217710

September 2012

AIAA–2012–3911

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

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Atlanta, Georgia, July 30–August 1, 2012
Acknowledgments

The work presented in this paper was supported by the Subsonic Fixed Wing Project at NASA Glenn Research Center. Special 

thanks to Jeff Berton and Bill Haller of NASA Glenn Research Center and Mark Guynn of NASA Langley Research Center 

for their support and guidance in this research. Lastly, the authors would like to thank General Electric for providing access to 

advanced counter-rotating propeller data in support of this study. 

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

This report contains preliminary findings, 

subject to revision as analysis proceeds.
NASA/TM—2012-217710 1

Performance and Weight Estimates for an 

Advanced Open Rotor Engine

Eric S. Hendricks and Michael T. Tong

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

NASA’s Environmentally Responsible Aviation Project and Subsonic Fixed Wing Project are 

focused on developing concepts and technologies which may enable dramatic reductions to the 

environmental impact of future generation subsonic aircraft. The open rotor concept (also historically 

referred to an unducted fan or advanced turboprop) may allow for the achievement of this objective by 

reducing engine fuel consumption. To evaluate the potential impact of open rotor engines, cycle modeling 

and engine weight estimation capabilities have been developed. The initial development of the cycle 

modeling capabilities in the Numerical Propulsion System Simulation (NPSS) tool was presented in a 

previous paper. Following that initial development, further advancements have been made to the cycle 

modeling and weight estimation capabilities for open rotor engines and are presented in this paper. The 

developed modeling capabilities are used to predict the performance of an advanced open rotor concept 

using modern counter-rotating propeller designs. Finally, performance and weight estimates for this 

engine are presented and compared to results from a previous NASA study of advanced geared and directdrive turbofans.

Introduction

At the outset of the Subsonic Fixed Wing (SFW) Project, NASA established a set of system level 

environmental metrics as a means to spur technology development to reduce aviation’s environmental 

impact. The goals, as shown in Figure 1, are generational and the metrics, and their respective values, 

have evolved over time and are relative to a particular baseline vehicle. The current metrics address 

reducing certification noise, oxides of nitrogen (NOx) emissions in the landing/takeoff (LTO) cycle, NOx 

emissions during cruise, and aircraft fuel/energy consumption (Ref. 1). The SFW project, NASA’s

Environmentally Responsible Aviation (ERA) Project, and the FAA’s Continuous Lower Energy, 

Emissions and Noise (CLEEN) program are all developing advanced technologies for reducing the 

environmental impact of subsonic transport aircraft across the range of future generations. While there is 

a desire to improve all subsonic aircraft, it is recognized that the benefits of future technologies depend on 

when those technologies mature to a sufficient Technology Readiness Level as well as on the vehicle size 

and mission.

To assess propulsion technologies capable of meeting the first generation of goals (N+1), the SFW 

systems analysis team has been investigating potential environmental benefits of advanced propulsion 

systems on an “advanced” single-aisle transport aircraft (ASAT). This aircraft would fall into the same 

class as the Boeing 737 and Airbus A320. Nominally, the ASAT aircraft would be capable of carrying 

162 passengers on a 3250 n mi flight. To power this aircraft, three types of advanced propulsion systems 

have been considered in a multiphase, multiyear study conducted by the systems analysis team. During 

the first 3 years of the study, two types of ultra-high bypass (UHB) turbofan concepts were assessed.

These two concept types were the advanced direct-drive turbofan and the geared turbofan. The studies 

showed potential reductions of 29 percent in fuel burn, 50 percent in LTO NOx emissions and a 

25 EPNdB reduction in cumulative noise relative to a late 1990s technology baseline (Ref. 2).
NASA/TM—2012-217710 2

Figure 1.—NASA's technology goals for subsonic transport aircraft (Ref. 1).

Figure 2.—A scale model of the open rotor being tested in 

NASA Glenn Research Center's 9- by 15-Ft low speed 

wind tunnel.

Following several phases of evaluating UHB turbofan concepts, the systems analysis team began 

investigating the open rotor concept for powering the ASAT aircraft. The open rotor engine concept, 

shown being tested in Figure 2, was originally developed during the 1970s and 1980s in response to high 

fuel prices. During that time, NASA’s Advanced Turboprop Project (ATP) sponsored research by

government, academia and industry to develop advanced propeller configurations. While the technology 

developed during the ATP never became commercial product, it demonstrated the potential to reduce 

aircraft fuel consumption by 25 to 30 percent versus aircraft flying at that time (Ref. 3). As a result, the 

ATP team (both NASA and industry partners) received the 1987 Collier Trophy “for developing 

advanced turboprop propulsion technology for new fuel efficient subsonic aircraft propulsion systems”
NASA/TM—2012-217710 3

(Ref. 3). With the return of higher fuel costs, the open rotor concept has received renewed interest due to 

its potential to reduce fuel consumption. Research efforts on developing and testing new open rotor blade 

designs are ongoing in the United States (Ref. 4) and in Europe(Ref. 5).

In order to evaluate the potential benefits of a modern open rotor engine, the SFW systems analysis 

team has developed modeling capabilities for assessing this system. Initial capabilities for analyzing the 

engine performance, aircraft mission and acoustic characteristics have been created for open rotors (Ref .

6 and 7). This paper first presents an overview of the previous cycle modeling developments made for 

open rotor engines. It then describes recent enhancements made to the cycle modeling capability as well 

as the development of open rotor engine weight estimation techniques. Next, the cycle and weight 

modeling techniques described in the paper are used to evaluate an open rotor engine for the ASAT 

aircraft. Finally, the open rotor engine developed for this application is compared to advanced geared and 

direct-drive turbofans developed during previous phases of the advanced propulsion study for ASAT 

aircraft.

Review of Previous Open Rotor Cycle Modeling Developments

Over the past 3 years, NASA has worked to develop cycle modeling capabilities for open rotor 

engines within the Numerical Propulsion System Simulation (NPSS) tool (Ref. 8). These modeling 

improvement efforts have concentrated in four main areas and were presented in a previous paper by the 

author (Ref. 6). The following sections briefly summarize the relevant modeling developments produced 

during the first 2 years of this research effort.

Counter-Rotating Propeller Modeling

The first challenge encountered when creating a cycle model for an open rotor engine was developing 

a performance prediction modeling capability for the counter-rotating propellers as the traditional 

propeller element in NPSS was not configured to handle the high speed, counter-rotating propellers. The 

traditional propeller element could not accept the multiple shaft inputs for the counter-rotating 

configuration and could not accept performance maps that were a function of freestream Mach number.

Therefore, a new component was created for NPSS which would enable analysis of this type of 

configuration. This component was written in a similar fashion as other NPSS components, with the 

performance of the counter-rotating propellers being estimated via performance maps. As a result, a wide 

variety of propeller designs can be modeled by simply replacing the performance map.

The development of this performance analysis capability in NPSS was based on the available 

information for some of the historic blade sets designed during the ATP. The most complete data set 

available from the ATP was for the F7/A7 blade geometry (Ref. 9). This scale model blade set was 

thoroughly tested in wind tunnels and was eventually used in tests of a full scale engine, the GE36 

Unducted Fan (UDF) Demonstrator. Using the available data, performance maps were created by fitting 

the data with a set of second-order response surface equations. These equations captured the relationship 

between the independent parameters (advance ratio, free stream Mach number and blade pitch angle) and 

the output power coefficient and thrust coefficient.

Counter-Rotating Turbine Modeling

During the initial development of the open rotor cycle analysis capability, a power turbine similar to 

that used in the GE36 UDF Demonstrator was assumed as the most likely configuration. In this 

architecture, an alternating blade row, counter-rotating turbine was used as shown in Figure 3 (Ref. 10).

This type of turbine is unique and required the investigation of modeling techniques that may be able to 

capture the performance of this type of system. Of the techniques tested, the use of a single traditional 

turbine component in NPSS was selected as it produced a reasonable representation based on available 

data from the GE36.

s
NASA/TM—2012-217710 4

Figure 3.—Schematic of GE36 power turbine (Ref. 10).

Multidesign Point (MDP) Approach

One of the key components of the initial open rotor modeling development was the implementation of 

a MDP approach for engine cycle analysis. In this approach, the NPSS engine model was set up such that 

several operating conditions would be simultaneously used to size the engine. This approach requires 

some additional implementation time, but results in all engines created during the design process meeting 

the operating requirements at all flight conditions. During the open rotor model development process, two 

design points were selected as being critical for the engine sizing process. These two points were a top-ofclimb condition (nominally Mach 0.78 at 35,000 ft, standard day) and a hot day, rolling takeoff condition 

(Mach 0.25, sea level, International Standard Atmosphere (ISA) +27 °F). By providing thrust targets and 

operating limits at these two points, an engine cycle can be quickly generated which meets these 

operating requirements for a given set of component designs and technology level assumptions. Through 

implementing the MDP approach, the development time for generating engine cycles that meet the 

aircraft mission needs was greatly reduced.

Power Management Strategy

Finally, a power management strategy was developed and implemented for the open rotor engine. The 

power management strategy is important for an open rotor as both the combustor fuel flow and the 

propeller blade pitch angles could be changed to affect thrust and fuel burn of the engine. From the 

historic development of the open rotor concept during the Advanced Turboprop Project, very little 

information was provided about how the engine would be controlled with these two degrees of freedom.

Therefore, a control strategy was created during the initial model development in which the blade pitch 

was varied with freestream Mach number according to a fixed schedule as shown in Figure 4 (Ref. 6).

This schedule was selected based on observing the impact of several pitch angles on the engine operation 

at various flight conditions. By selecting this power management strategy, the engine could easily be 

evaluated at all operating conditions in the flight envelope. 
NASA/TM—2012-217710 5

Figure 4.—Front propeller blade angle schedule (Ref. 6).

Open Rotor Modeling Improvements

The initial modeling developments presented in the previous sections allowed for the cycle analysis 

of a notional open rotor engine to be completed. However, over the past year a number of improvements 

have been made to NASA open rotor modeling capabilities which have enhanced the ability to assess 

future open rotor engines. The following sections briefly summarize the recent modeling developments 

made for the open rotor which supplement or replace the modeling capabilities described above.

Performance Maps for Modern Counter-Rotating Propellers

In the initial development of an open rotor engine cycle model described in Reference 6, the historic 

F7/A7 counter-rotating propellers were used to generate the performance maps. These propellers were 

designed during the 1980s as part of the Advanced Turboprop Project and therefore do not accurately 

estimate performance characteristics of modern blade designs. As a result, new performance maps 

representative of modern blade designs were needed in order to estimate the benefits of a future open 

rotor engine.

Over the past few years, NASA and General Electric have been collaborating to design and test 

advanced open rotor blade geometries (Ref. 11). A multiyear test campaign was conducted in the NASA 

Glenn Research Center 8- by 6-Ft and 9- by 15-Ft wind tunnel complex that measured both performance 

and acoustic characteristics of these modern blade designs. From the performance characteristics 

observed in these tests, performance maps were created following the process described in Reference 6.

The performance maps correlate the thrust and power coefficients to advance ratio, free stream Mach 

number and blade pitch angle using a second order polynomial response surface equation. Using these 

performance maps in the cycle modeling has enabled the assessment of an advanced open rotor engine 

system.
NASA/TM—2012-217710 6

Figure 5.—Geared open rotor engine components.

Geared Open Rotor Architecture

In the initial development of the open rotor modeling capability, a direct drive architecture with a 

counter-rotating power turbine was assumed. This configuration eliminates the need for a large, heavy 

and complex gearbox. However, the direct drive configuration requires that the turbine and propellers 

rotate at the same speeds. The propellers designed for open rotors are large in diameter and constrained by 

tip speed, typically resulting in rotational speeds of less than 1100 rpm. This relatively slow rotation 

speed makes designing the turbine difficult as a larger diameter or more stages need to be added to extract 

enough work from the core flow. Larger diameters and additional stages add more weight to the turbine 

and the overall engine system. Alternatively, an engine with a gearbox allows the propeller and turbine to 

each be designed for different rotational speeds. The turbine can be designed to provide the same power at 

a higher rotation speed with a smaller diameter and fewer stages. This turbine design would be lighter 

than a counter-rotating turbine meeting the same power requirements and would offset the weight added 

by the gearbox. In a trade study comparing the geared and direct drive open rotor engines, similar 

performance levels were achieved with the weight estimate for geared engine slightly lower than the 

direct drive configuration. This benefit would likely be greater as the open rotor engine grows to higher 

thrusts. The higher thrust engines would need to have a larger propeller diameter to produce the required 

thrust and the tip speed limit would result in even slower propeller rotation speeds. Therefore, a cycle 

model for a geared open rotor engine was pursued for this study as shown in Figure 5. In this 

configuration, there are two spools composing the gas generator: the low pressure (LP) spool composed 

of a compressor (LPC) and turbine (LPT) and the high pressure (HP) spool composed of a compressor 

(HPC) and turbine (HPT). Downstream of the gas generator is a power turbine (PT) which drives the 

counter-rotating propellers through a gearbox. 

Power Management Strategy

During the initial development of the open rotor cycle modeling capability, a power management 

strategy was developed which utilized a blade pitch schedule as a function of Mach number. This 

schedule was required as both the propeller blade pitch angles and fuel flow could be varied to achieve 

the desired thrust. However, the strategy of holding blade pitch constant at a given Mach number resulted 

in large changes in propeller speed as the engine thrust was reduced. By comparison, most modern 

turboprops operate with constant speed propellers. Therefore, a constant tip speed power management 

strategy was investigated and implemented in the current open rotor model. The constant tip speed power 

management strategy was implemented in the NPSS cycle model by adding an independent parameter for 

the blade pitch and a dependent parameter for the tip speed to the solver. Additionally, a minimum blade 
NASA/TM—2012-217710 7

Figure 6.—Blade angle settings for constant tip speed throughout the flight envelope.

angle constraint was applied to prevent the engine from operating at blade angles that were outside the 

range of validity for the propeller maps. As a result of utilizing the constant tip speed strategy, operating 

points for the engine were identified for selected flight conditions as shown in Figure 6. In comparison to 

the previous power management strategy shown in Figure 4, the constant tip speed strategy shown in 

Figure 6 gives a range of possible blade angles for each flight Mach number. However, the overall trend 

of blade pitch angle increasing as Mach number increases remains similar to the previous strategy. The 

new constant tip speed approach also captures other minor factors that impact the blade angle, such as the 

variations in altitude. Figure 6 can provide further insight into how the blade angles change as the engine 

thrust is changed at a given flight condition. For the maximum thrust (power) conditions, the blade angle 

is the largest with the blade angle decreasing by as much as 10° as the engine throttle is reduced. 

Open Rotor Weight Estimation

Finally, the NASA software tool WATE++ (Weight Analysis of Gas Turbine Engines) (Refs. 12, 13,

and 14) was used to create an engine component architecture that would match the engine thermodynamic 

cycle produced using the methods in the previous sections. The cycle data required for WATE execution, 

such as airflow, temperatures, and pressures, pressure ratios, etc., were derived from the NPSS cycle 

model output. The maximum temperature and pressure required for sizing each engine component were 

identified by predicting the cycle performance throughout the flight envelope. The cycle data, the material 

properties, and design rules for geometric, stress, and turbomachinery stage-loading limits were used to 

determine an acceptable engine layout. The material assumed for the LPC and for the first five stages of 

the HPC is titanium. For the remaining stages of the HPC, a titanium-aluminide alloy is used. Nickelbased alloy is used for all the turbine components. For the counter-rotating propellers, a rotor weight 

correlation which is a function of maximum power delivered to the rotor, blade tip-speed, and number of 

blades is used. This correlation assumes the blades are constructed of polymer composites and is 

described in detail in Reference 15. An empirical correlation was used to calculate the weight of the 

gearbox and lubrication system as shown in Figure 7. The correlation is a function of maximum delivered 

output power and gear ratio, and was developed at NASA based on actual gearbox weight data from over 

fifty rotorcraft, tiltrotors, and turboprop aircraft.
NASA/TM—2012-217710 8

Figure 7.—Gearbox and Lubrication System Weight Correlation.

Open Rotor Modeling Results

The modeling developments described in the previous sections were used to create an open rotor 

engine model representative of a future system capable of powering an advanced single aisle transport 

aircraft. This aircraft is designed to fly 3250 n mi at a cruise Mach number of 0.78 carrying 162 

passengers. The following sections provide an overview of the performance and weight estimates 

calculated for an open rotor engine based on the described model. Following these sections, the open rotor 

engine results are compared to advanced geared and direct-drive turbofans which were modeled by 

NASA as part of a previous study.

Cycle Analysis and Performance Results

The cycle model was developed using the MDP strategy previously summarized in which two 

operating conditions were simultaneously used to size the engine components. The two design points 

selected for this study were a top-of-climb (Mach 0.78 at 35,000 ft) condition and a rolling takeoff (Mach 

0.25 at 0 ft, ISA+27 °F) condition. At each of these points, thrust requirements and operating limits were 

specified for the engine. Additionally, performance assumptions were specified for each core component 

at the top-of-climb (aerodynamic design point) flight condition as shown in Table 1. The component 

efficiency assumptions listed in the table are consistent assumptions used in a recent NASA study of 

advanced geared and direct drive turbofans (Ref. 2). However, the pressure ratios of the low pressure and 

high pressure compressors were adjusted to match the open rotor engine architecture. The LPC pressure 

ratio was set to 4.2 and the HPC pressure ratio set to 10.0 to give an overall pressure ratio of 42.0 which 

matches the turbofan study. In addition to the assumptions in Table 1, a gear ratio of 6 and a gearbox 

efficiency of 99 percent were assumed. Finally, propeller performance assumptions were selected based 

the propeller design characteristics and data obtained during wind tunnel testing.
NASA/TM—2012-217710 9

TABLE 1.—OPEN ROTOR ENGINE COMPONENT PERFORMANCE 

ASSUMPTIONS AT TOP OF CLIMB (M0.78, 35000 ft)

Component Parameter Value

LPC Pressure ratio 

Adiabatic efficiency, % 

4.2 

89.6 

HPC Pressure ratio 

Adiabatic efficiency, %

10.0 

88.6 

HPT Adiabatic efficiency, % 91.9 

LPT Adiabatic efficiency, % 94.2 

Power turbine Adiabatic efficiency, % 94.0 

Using MDP methodology and assumptions described above an NPSS cycle model was developed.

Table 2 provides a summary of key engine metrics at the two operating points used to design the engine 

cycle model as well as at sea level static. At the top-of-climb point, 5000 lb of thrust is required by the 

airplane to achieve service ceiling performance with an overall pressure ratio (OPR) of 42 coming from 

the component performance assumptions. At the rolling takeoff design point, 17,500 lb of thrust was

required to meet field length performance for the aircraft. However, 19,000 lb force of thrust was 

produced by the engine which exceeded the thrust required by 1,500 lb. The engine exceeded the rolling 

takeoff thrust required as a result of temperature constraints set in the engine MDP process. These 

constraints limited the combustor exit temperature at the top-of-climb flight condition to ensure the hot 

section would operate at lower temperatures during cruise resulting in longer life of the turbines. As a 

result, the open rotor engine produces excess thrust at the rolling takeoff and sea level static conditions.

TABLE 2.—OPEN ROTOR ENGINE PERFORMANCE FOR MAXIMUM THRUST AT KEY OPERATING CONDITIONS

Performance parameter Top of climb, 

M0.78, 35000 ft

Rolling takeoff, 

M0.25, 0 ft, ISA+27 °F

Sea level static, 

M0.0, 0 ft, ISA+27 °F

Net thrust, lbf 5000 19000 27300

Propeller thrust, lbf 4830 18600 26500

TSFC, lbm/hr/lbf 0.428 0.229 0.158

OPR 42.0 28.5 29.4

Combustor exit temperature, R 3280 3460 3460

Core mass flow, lbm/s 21.8 42.7 42.2

For all three flight conditions shown in Table 2, a low thrust specific fuel consumption (TSFC) is 

predicted from the cycle model. The low TSFC is a consequence of having a large portion of the net 

thrust at each operating condition being produced by the high-efficiency counter-rotating propellers. The 

low TSFC is a highly desirable characteristic of the open rotor engines and persists throughout the flight 

envelope. The TSFC characteristics for different throttle settings at the flight conditions of Table 2 are 

shown in Figure 8.

Weight and Flowpath Estimates

Following the engine cycle model development, an estimate of the engine weight and flowpath 

dimensions was developed using the methods described in the previous section. Table 3 gives a summary 

of the major weights and dimensions of the open rotor engine developed for this study. The overall engine 

weight is predicted to be over 9,000 lb, with about a third of the engine weight coming from the counterrotating propellers. An additional 1000 lb of the engine weight comes from the addition of a gearbox to 

the engine. Table 3 also lists the estimated overall dimensions of the open rotor engine. The maximum 

propeller diameter for the engine designed in this study is estimated to be about 13.8 ft with an overall 

engine length of about 23 ft. Additionally, Table 3 presents the estimates for the number of stages for 

each turbomachinery component and the assumed gear ratio for the gearbox. A notional sketch of the 
NASA/TM—2012-217710 10

flowpath of the open rotor engine developed in this study is shown in Figure 9. The center box between 

the propeller blades is the location of the gearbox, with the other boxes representing ducts that pass the 

core flow around the turbine to the core nozzle.

Figure 8. Power Hooks for Top of Climb, Rolling Takeoff and Sea Level Static Operating Conditions.

TABLE 3.—PRINCIPLE MECHANICAL DESIGN PARAMETERS

Mechanical design parameter Value

Total engine pod weight, lbm ................................................................9220

Counter-rotating propeller weight, lbm .................................................3240

Gearbox weight, lbm .............................................................................1030

Gas generator, turbomachinery weight, lbm..........................................4950

Maximum propeller diameter, ft.............................................................13.8

Nacelle diameter, ft...................................................................................5.6

Overall length, ft.....................................................................................23.2

LPC stages...................................................................................................4

HPC stages...................................................................................................7

HPT stages...................................................................................................1

LPT stages...................................................................................................1

Power turbine stages....................................................................................3

Gear ratio.....................................................................................................6
NASA/TM—2012-217710 11

Figure 9.—Notional Cross-Section of the Open Rotor Engine.

Comparison to NASA Advanced Turbofan Study

The open rotor engine study described in this paper is the latest in a series of NASA studies which 

have examined future propulsion system concepts for an ASAT aircraft. As such, the open rotor model 

development followed a similar modeling process and applied the same technology assumptions as the 

previous studies. Therefore, a comparison can be made between engines of the previous studies and the 

open rotor engine presented in this paper.

The previous studies on advanced propulsion systems for ASAT aircraft focused on modeling geared 

and direct drive turbofans as described in Reference 2. In the most recent phase of the turbofan study, two 

engines, one geared and one direct drive, were identified as desirable due to their system level 

characteristics. A geared turbofan with a high pressure ratio LPC and a design point fan pressure ratio of 

1.5 was determined to be the engine that resulted in the lowest block fuel (and therefore the lowest 

operating cost) required for the selected mission. The second engine identified in the study was a direct 

drive configuration with a low pressure ratio LPC and design point fan pressure ratio of 1.7. This engine 

resulted in the lowest aircraft takeoff gross weight which is often used as a surrogate parameter in 

conceptual design for the purchase cost of the aircraft. Because of the desirable features of these two 

engines, they were selected for comparison to the open rotor engine described in this paper.

Key engine characteristics for the geared turbofan, direct drive turbofan and open rotor engines are 

presented in Table 4. At the top-of-climb flight condition, all three engines produce a similar amount of 

thrust (see table note regarding differences in top-of-climb flight conditions between the turbofans and 

open rotor) with the same overall pressure ratio. While all three engines produce similar amounts of 

thrust, the open rotor engine has a lower thrust specific fuel consumption. A lower TSFC generally 

translates to less fuel required to fly the same mission with the aircraft. At the sea level static operating 

point, the open rotor engine performance and environmental characteristics are noticeably different than

both of the turbofan engines. First, the open rotor engine produces almost 4000 lb of additional thrust.

The higher thrust at sea level is again evidence that the open rotor engine is oversized for takeoff because 

of the requirements and constraints placed on the engine at the top-of-climb design point. The open rotor 

engine also has a lower TSFC at the sea level static flight condition compared to both turbofans, 

indicating lower fuel consumption for the open rotor engine during takeoff.
NASA/TM—2012-217710 12

TABLE 4.—COMPARISON OF ADVANCED TURBOFAN AND OPEN ROTOR ENGINES

Engine parameter Geared 

turbofan

Direct drive 

turbofan

Open rotora

Top of climb,

M0.80, 35000 ft

OPR 42 42 42

Net thrust, lbf 5000 5000 4960

TSFC, lbm/hr/lbf 0.502 0.525 0.441

Sea level static, 

M0.0, 0 ft, ISA+27 °F

OPR 32.7 33.8 29.4

Net thrust, lbf 23400 22700 27300

TSFC, lbm/hr/lbf 0.257 0.290 0.158

Fan or propeller diameter, ft 6.2 5.5 13.8

Nacelle maximum diameter, ft 7.6 6.7 5.6

Total engine pod weight, lbm 6630 6100 9220 a

Note that the open rotor engine was designed for a top of climb flight condition of M0.78 at 

35000 ft in order to match propeller design characteristics. This table presents open rotor engine 

performance characteristics at M0.80, 35000 ft for top of climb for comparison to the previously 

designed turbofan engines of Reference 2.

Table 4 also provides a comparison of some of the major mechanical design characteristics of the 

three engines. The open rotor engine is much larger in diameter as the propellers are almost twice the

diameter of the turbofan nacelles. Furthermore, the open rotor engine weight is significantly heavier, 

primarily due to the large propellers. The weight and size of the open rotor will partially offset the benefit 

of lower fuel burn for this engine compared to the turbofans. An aircraft with an open rotor engine will 

likely have a higher empty weight (due to the higher engine weight), but the aircraft would require less 

fuel to complete the mission. As a result, the overall aircraft takeoff gross weight would likely be similar 

between turbofan and open rotor powered aircraft (Ref. 16).

Conclusion

NASA’s Subsonic Fixed Wing and Environmentally Responsible Aviation Programs are both 

interested in developing technologies which will reduce the environmental impact of future aircraft. The 

open rotor engine is of interest to both of these programs due to its potential to reduce fuel consumption 

and emissions of NOx near airports. In order to assess the potential impact of an open rotor engine on an 

advanced single aisle aircraft, modeling capabilities have been developed by NASA to predict the 

performance and weight characteristics of open rotor engines. The initial development of open rotor 

modeling capabilities was described in a previous paper.

Following the initial model development, further improvements to open rotor modeling capabilities 

were made and presented in this paper. First, propeller maps representative of modern blade designs were 

developed using data collected in recent wind tunnel experiments by NASA and General Electric. Second, 

the model was transitioned to a geared open rotor architecture as the geared configuration provided better 

or equivalent performance at a lower overall engine weight. Next, a new power management strategy was 

implemented to control the blade pitch angles throughout the flight envelope. The new constant tip speed 

strategy replaced the assumed blade angle schedule and provided a more realistic approximation of how 

an open rotor engine would be operated. Lastly, weight estimation techniques were developed for the 

counter-rotating propellers and the gearbox. These methods were then implemented to determine the 

mechanical design characteristics of an open rotor engine.

Using the modeling capabilities summarized in this paper, an open rotor engine model was created.

This engine was sized using a MDP approach to power an advanced single aisle transport aircraft. The 

open rotor engine model predicted a low thrust specific fuel consumption throughout the flight envelope.

In addition, a large thrust lapse resulted from the MDP design process producing an engine that is 

oversized at rolling takeoff and sea level static flight conditions. The open rotor weight was estimated to 

be 9220 lb, with about half of the engine pod weight coming from the propellers and gearbox. Lastly, the 

open rotor model results were compared to similar result for both geared and direct drive turbofans. The 
NASA/TM—2012-217710 13

open rotor system is predicted to have better thrust specific fuel consumption compared to both turbofans.

However, the turbofan engines are both about 3000 lb lighter than the open rotor engine. 

While a direct comparison of the open rotor, geared turbofan and direct drive turbofan can provide 

some insight into the benefits of each, the impact of the engine can only be truly assessed when they are 

integrated with a full aircraft. By integrating the engine with an aircraft, total fuel burn, emissions, and 

noise estimates can be made. An assessment of an entire aircraft system for these metrics was beyond the 

scope of this paper. However, the open rotor engine described in this paper has been integrated with an 

advanced aircraft to evaluate the potential benefit of this engine. Overall, the open rotor powered aircraft 

is predicted to have 36 percent lower fuel consumption than a 1990s reference baseline aircraft (Ref. 16).

Additionally, the open rotor aircraft is estimated to have a 13 dB cumulative margin to Stage 4 noise 

regulations (Ref. 16). While there is uncertainty in both of these estimates, they indicate that an open rotor 

engine may be able to meet the future propulsion requirements of a single aisle transport aircraft. 

References

1. Del Rosario, R., “NASA Subsonic Fixed Wing Project Overview,” 2012 NASA Fundamental 

Aeronautics Program Technical Conference, Renaissance Cleveland Hotel, Cleveland, 14 March 

2012, Lecture.

2. Guynn, M., Berton, J., Fisher, K., Haller, W., Tong, M. and Thurman, D., “Refined Exploration of 

Turbofan Design Options for an Advanced Single-Aisle Transport,” NASA/TM—2011-216883, 

2011.

3. Hager, Roy D., and Vrabel, Deborah, “Advanced Turboprop Project”, NASA SP–495, 1988. 

4. Van Zante, D., “Open Rotor Noise Research in the US.” X-Noise Workshop, Lausanne, Switzerland, 

18 March 2011, Lecture.

5. Parry, A. B., “Current Open Rotor Noise Research in EU,” X-Noise Workshop, Lausanne, 

Switzerland, 18 March 2011, Lecture.

6. Hendricks, E. “Development of an Open Rotor Cycle Model in NPSS Using a Multi-Design Point 

Approach.” Presented at the 2011 ASME Turbo Expo in Vancouver, Canada. GT2011-46694. 

7. Guynn, M.; Berton, J.; Hendricks, E.; Tong, M.; Haller, W.; and Thurman, D., “Initial Assessment of 

Open Rotor Propulsion Applied to an Advanced Single-Aisle Aircraft,” AIAA-2011-7058, 10th 

AIAA Aviation Technology, Integration, and Operations (ATIO) Conference, September, 2011, 

Virginia Beach, VA, 20-22 September, 2011. 

8. NPSS User Guide Software Release: NPSS_1.6.5.

9. Hoff, G.E. et al., “Experimental Performance and Acoustic Investigation of Modern, Counterrotating 

Blade Concepts,” NASA-CR-185158, 1990.

10. Reid, C. “Overview of Flight Testing of GE Aircraft Engines UDF Engine.” AIAA-88-3082.

11. Norris, G., “Open Return: General Electric and NASA dust off test rigs from 20 years ago,” Aviation 

Week and Space Technology, 14 July 2008.

12. Onat, E. and Klees, G.W., “A Method to Estimate Weight and Dimensions of Large and Small Gas 

Turbine Engines,” NASA CR–159481, 1979.

13. Tong, M.T., Halliwell, I., Ghosn, L.J., “A Computer Code for Gas Turbine Engine Weight and Life 

Estimation,” ASME Journal of Engineering for Gas Turbine and Power, volume 126, no. 2, pp.265-

270, April 2004.

14. 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.

15. Weisbnch, A. L., Godston, J. and Bradle, E.S., “Technology and Benefits of High Speed Aircraft 

Counter Rotation Propellers,” NASA CR–168258, 1982.

16. Guynn, M., “Open Rotor Propulsion Assessment for Advanced Single-Aisle Aircraft,” 2012 NASA 

Fundamental Aeronautics Program Technical Conference, Renaissance Cleveland Hotel, Cleveland, 

14 March 2012, Lecture.
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Performance and Weight Estimates for an Advanced Open Rotor Engine 

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14. ABSTRACT

NASA’s Environmentally Responsible Aviation Project and Subsonic Fixed Wing Project are focused on developing concepts and 

technologies which may enable dramatic reductions to the environmental impact of future generation subsonic aircraft. The open rotor 

concept (also historically referred to an unducted fan or advanced turboprop) may allow for the achievement of this objective by reducing 

engine fuel consumption. To evaluate the potential impact of open rotor engines, cycle modeling and engine weight estimation capabilities 

have been developed. The initial development of the cycle modeling capabilities in the Numerical Propulsion System Simulation (NPSS) 

tool was presented in a previous paper. Following that initial development, further advancements have been made to the cycle modeling and 

weight estimation capabilities for open rotor engines and are presented in this paper. The developed modeling capabilities are used to predict 

the performance of an advanced open rotor concept using modern counter-rotating propeller designs. Finally, performance and weight 

estimates for this engine are presented and compared to results from a previous NASA study of advanced geared and direct-drive turbofans. 

15. SUBJECT TERMS

Propellers; Contrarotating propellers; Turboprop engines; Gas turbine engines 

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