October 2012 

NASA/TM-2012-217772 

Performance and Environmental Assessment of 

an Advanced Aircraft with Open Rotor 

Propulsion 

Mark D. Guynn 

Langley Research Center, Hampton, Virginia 

Jeffrey J. Berton, William J. Haller, Eric S. Hendricks, and Michael T. Tong 

Glenn Research Center, Cleveland, Ohio 
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October 2012 

NASA/TM-2012-217772 

Performance and Environmental Assessment of 

an Advanced Aircraft with Open Rotor 

Propulsion 

Mark D. Guynn 

Langley Research Center, Hampton, Virginia 

Jeffrey J. Berton, William J. Haller, Eric S. Hendricks, and Michael T. Tong 

Glenn Research Center, Cleveland, Ohio 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320

443-757-5802 

Acknowledgments 

This study was supported by the Subsonic Fixed Wing project of NASA’s Fundamental 

Aeronautics Program. Experimental data used in this study were obtained with support 

from the Environmentally Responsible Aviation project of NASA’s Integrated Systems 

Research Program. Both this study and the open rotor wind tunnel test campaign were 

conducted in a collaborative partnership with General Electric Aviation. 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration.
iii

Contents 

List of Tables ............................................................................................................................................... iv

List of Figures ............................................................................................................................................... v

List of Acronyms, Symbols, and Abbreviations .......................................................................................... vi

Abstract ......................................................................................................................................................... 1

1.0 Introduction ............................................................................................................................................. 1

2.0 Study Objectives and Approach .............................................................................................................. 3

3.0 Modeling and Analysis Methodology ..................................................................................................... 3

3.1 Propulsion Modeling ........................................................................................................................ 3

3.2 Airframe Modeling ........................................................................................................................... 4

3.3 Noise Analysis .................................................................................................................................. 7

3.3.1 Open Rotor Source Noise Modeling .................................................................................... 7

3.3.2 System-Level Noise Predictions ........................................................................................ 10

3.3.3 Special Considerations for Open Rotor Aircraft ................................................................ 11

3.3.4 Establishing Reference System Noise Levels .................................................................... 12

3.3.5 System Corrections to Reference Noise Levels ................................................................. 14

4.0 Analysis Results .................................................................................................................................... 15

4.1 Engine Design ................................................................................................................................ 15

4.2 Aircraft Sizing and Performance .................................................................................................... 17

4.3 Certification Noise ......................................................................................................................... 20

4.3.1 Trajectory Analysis ............................................................................................................ 20

4.3.2 Noise at Observers ............................................................................................................. 22

5.0 Concluding Remarks ............................................................................................................................. 23

6.0 References .......................................................................................................... 25

Appendix................................................................................................................. 28
iv

List of Tables 

Table 1. General Engine Characteristics ..................................................................................................... 16

Table 2. Aircraft Sizing Results .................................................................................................................. 18

Table 3. Fuel Consumption (lb) by Mission Segment ................................................................................ 20

Table 4. Effective Perceived Noise Levels (EPNdB) for a 151,300 lb Aircraft ......................................... 23
v 

List of Figures 

Figure 1. NASA’s technology goals for subsonic transport aircraft. ............................................................ 2

Figure 2. Rear-mounted open rotor pusher configuration, shown on Boeing 727 test vehicle. .................... 4

Figure 3. Process used to model advanced technology airframe suitable for open rotor propulsion. ........... 5

Figure 4. Basic Advanced Single-Aisle Transport configuration with rear-mounted open rotor engines. ... 7

Figure 5. Open rotor test article and linear microphone array in the NASA Glenn 9x15-ft Low Speed 

Wind Tunnel. ................................................................................................................................................ 8

Figure 6. Sample open rotor narrowband spectrum conversions. ............................................................... 10

Figure 7. Noise observer arrangement relative to takeoff and landing aircraft trajectories. ....................... 11

Figure 8. Example flyover observer PNLT noise histories. ........................................................................ 13

Figure 9. Advanced engine architectures. ................................................................................................... 17

Figure 10. Advanced vehicle weight, fuel, and emissions relative to 1990s technology baseline. ............. 18

Figure 11. Block fuel comparison for advanced configurations. ................................................................ 20

Figure 12. Takeoff and approach trajectory, aircraft parameters. ............................................................... 21

Figure 13. Takeoff and approach trajectory, engine parameters. ................................................................ 22

Figure 14. Reference system EPNL for the flyover observer. .................................................................... 23
vi

List of Acronyms, Symbols, and Abbreviations 

cFlight – speed of sound at aircraft flight conditions 

cTunnel – speed of sound at wind tunnel test conditions 

Dp – mass of nitrogen oxides emitted over a standard landing-takeoff cycle

DFull-Scale – diameter of rotor for full scale engine 

DRig – diameter of rotor for wind tunnel model 

fFlight – frequency of rotor noise at aircraft flight conditions 

fStatic – frequency of rotor noise at static conditions 

fTunnel – frequency of rotor noise at wind tunnel test conditions 

Foo – engine rated output at sea-level, static conditions 

M – Mach number

MFlight – flight Mach number 

MTunnel – wind tunnel test Mach number 

SME – source motion exponent 

 – aircraft angle of attack 

Cant – rotor cant angle relative to aircraft zero angle of attack reference line 

Inflow – rotor inflow angle 

 – downwash angle at rotor location (angle of flow relative to freestream) 

  – downwash angle at rotor location with aircraft angle of attack of zero 

G – geometric sideline microphone array angles 

E – noise emission yaw (polar) angles 

Flight – air density at aircraft flight conditions 

Tunnel – air density at wind tunnel test conditions 

ACD – Acoustics Data Module 

AFE – Above Field Elevation 

ANOPP – Aircraft Noise Prediction Program 

ASAT – Advanced Single-Aisle Transport 

ASAT-or – Advanced Single-Aisle Transport-open rotor 

ASAT-re – Advanced Single-Aisle Transport-rear engine (turbofan) 

CSAT-re – Current technology Single-Aisle Transport-rear engine (turbofan) 

EI – Emission Index, grams of emission per kilogram of fuel consumed 

EPNL – Effective Perceived Noise Level 

ERA – NASA Environmentally Responsible Aviation project 

FAA – Federal Aviation Administration 
vii 

FAR – Federal Aviation Regulations 

FLOPS – Flight Optimization System 

FPR – Fan Pressure Ratio 

ICAO – International Civil Aviation Organization 

ISA – International Standard Atmosphere 

L/D – Lift-to-Drag ratio 

LTO – Landing-Takeoff Cycle 

NASA – National Aeronautics and Space Administration 

NOX – Nitrogen Oxides (NO and NO2) 

NPSS – Numerical Propulsion System Simulation 

OEW – Operating Empty Weight 

PDCYL – Point Design of Cylindrical-bodied aircraft 

PNL – Perceived Noise Level 

PNLT – Perceived Noise Level, tone-corrected 

SFW – NASA Subsonic Fixed Wing project 

SL – Sea Level 

SLS – Sea Level, Static 

SPL – Sound Pressure Level 

TOC – Top-Of-Climb 

TSFC – Thrust Specific Fuel Consumption 

T/W – Thrust-to-Weight ratio 

WATE – Weight Analysis of Turbine Engines 
1 

Abstract 

Application of high speed, advanced turboprops, or “propfans,” to 

transonic transport aircraft received significant attention during the 

1970s and 1980s when fuel efficiency was the driving focus of 

aeronautical research. Unfortunately, after fuel prices declined sharply 

there was no longer sufficient motivation to continue maturing this 

technology. Recent volatility in fuel prices and increasing concern for 

aviation’s environmental impact, however, have renewed interest in 

unducted, open rotor propulsion and revived research by NASA and a 

number of engine manufacturers. Because of the renewed interest in 

open rotor propulsion, the lack of publicly available up-to-date studies 

assessing its benefits, and NASA’s focus on reducing fuel consumption, a 

preliminary aircraft system level study on open rotor propulsion was 

initiated to inform decisions concerning research in this area. New 

analysis processes were established to assess the characteristics of open 

rotor aircraft. These processes were then used to assess the performance, 

noise, and emissions characteristics of an advanced, single-aisle aircraft 

using open rotor propulsion. The results of this initial study indicate 

open rotor engines have the potential to provide significant reductions in 

fuel consumption and landing-takeoff cycle NOX emissions compared to 

aircraft utilizing turbofan engines with equivalent core technology. In 

addition, noise analysis of the study configuration indicates that an open 

rotor aircraft in the single-aisle class would be able to meet current 

noise regulations with margin. 

1.0 Introduction 

The Subsonic Fixed Wing (SFW) Project of NASA’s Fundamental Aeronautics Program and the 

Environmentally Responsible Aviation (ERA) Project of NASA’s Integrated System Research Program 

have jointly established a series of technology goals for future generations of subsonic transport aircraft. 

These goals are shown in Figure 1, where “N+x” refers to the series of technology generations which 

emerge over time. Propulsion technology will play a critical role in reaching the goals in every 

generation. Engines of the future will need to have lower fuel consumption, lower noise, and emit fewer 

harmful emissions. Airlines, and therefore aircraft and engine manufacturers, have always had a desire to 

reduce fuel consumption because of its direct impact on operating cost. But because of the significant 

risks inherent in adopting something completely new, a great deal of economic pressure is necessary for 

revolutionary technologies to be implemented. In NASA’s opinion, however, the goals in Figure 1 cannot 

be achieved with conservative, evolutionary technologies and concepts. As NASA works towards these 

goals by investing in technology research and development, numerous unconventional ideas are 

considered and evaluated for their potential impacts. One such unconventional concept is open rotor 

propulsion. 
2 

N+1 (2015) N+2 (2020**) N+3 (2025)

Noise

(cum margin rel. to Stage 4) -32 dB -42 dB -71 dB

LTO NOx Emissions

(rel. to CAEP 6) -60% -75% -80%

Cruise NOx Emissions

(rel. to 2005 best in class) -55% -70% -80%

Aircraft Fuel/Energy Consumption‡

(rel. to 2005 best in class)

-33% -50% -60%

‡ CO2 emission benefits dependent on life-cycle CO2e per MJ for fuel and/or energy source used

TECHNOLOGY GENERATIONS

(Technology Readiness Level = 4-6)

* Projected benefits once technologies are matured and implemented by industry. Benefits vary by vehicle size and mission. N+1 and N+3 values 

 are referenced to a 737-800 with CFM56-7B engines, N+2 values are referenced to a 777-200 with GE90 engines

** ERA's time-phased approach includes advancing "long-pole" technologies to TRL 6 by 2015

TECHNOLOGY

BENEFITS*

Figure 1. NASA’s technology goals for subsonic transport aircraft. 

When the aviation industry faced dramatically rising fuel costs in the 1970s, the response was a 

national research portfolio aimed at improving aircraft fuel efficiency (refs. 1, 2). One element of this 

research portfolio was the NASA Advanced Turboprop Project, focused on high-speed, unducted (or 

“open rotor”) propulsors (ref. 3). A number of open rotor concepts were developed and tested as part of 

this program. The program was highly successful and was awarded the Collier Trophy in 1987; but after 

fuel prices declined sharply there was no longer sufficient motivation to continue maturing the technology 

and the ducted turbofan engine continues to be the standard propulsion system for high-speed commercial 

transport aircraft. However, recent volatility in fuel prices and increasing concern for aviation’s 

environmental impact have renewed interest in unducted, open rotor propulsion for airliners and revived 

research by NASA and a number of engine manufacturers (refs. 4, 5).

Numerous studies on the benefits of open rotor engines were conducted in the 1970s and 1980s; 

references 6 and 7 describe two of many examples. Generally, the target applications in these studies were 

short range aircraft in the 100-150 passenger class with cruise Mach numbers less than 0.8. The fuel 

consumption benefit of open rotor propulsion compared to equivalent technology turbofan engines was 

typically found to be around 25%. One of the challenges consistently identified, however, was higher 

noise. Interior noise was reduced to levels equivalent to aircraft of that era through the use of additional 

fuselage acoustic treatment, at the cost of increased aircraft empty weight. Most of the studies found 

certification noise levels at or below FAA Stage 3/ICAO Chapter 3 limits (see references 8 and 9 for a 

description of aircraft noise certification). Although commercial transport aircraft designs with open rotor 

propulsion were studied extensively in the past, the results of those studies do not necessarily remain 

valid today due to advances in airframe and engine technology, and changes in the regulatory and market 

environments into which the aircraft will be inserted. Benefits measured relative to the engines of that era 

are no longer valid since ducted turbofan engines have advanced in the intervening years. Also, current 

and future noise regulations are stricter than those considered in the past studies. Because of the renewed 

interest in open rotor propulsion, the lack of publicly-available up-to-date studies assessing its benefits, 

and NASA’s focus on reducing fuel consumption, internal NASA studies on open rotor propulsion were 

initiated to inform decisions concerning research in this area. These system-level studies leverage an open 

rotor research partnership between General Electric Aviation and NASA’s Environmentally Responsible 

Aviation Project. 
3 

Unfortunately, in the two decades that have passed since open rotor concepts were thoroughly 

investigated, NASA has lost experience and expertise in this technology area. In order for an open rotor 

aircraft to be assessed to the same level of confidence as more traditional designs, new tools and methods 

are needed. The first step in performing a reliable assessment of this technology is to develop the 

necessary analysis processes. This report describes the initial capabilities which have been established in 

the SFW project to assess aircraft-level characteristics for designs with open rotor propulsion. Many of 

these open rotor specific analysis procedures were developed in cooperation with General Electric 

Aviation. The results of an aircraft system level assessment of open rotor propulsion applied to a Boeing 

737/Airbus 320 class vehicle are then presented and compared to those for an equivalent technology 

geared turbofan based design. A potential new vehicle in this size class would be a prime candidate for 

open rotor propulsion because the 737/A320 class represents a significant portion of the global airline 

fleet, and a significant opportunity for technology insertion. Sixty-nine percent of the new aircraft 

produced over the next 20 years are projected to be in this class (ref. 10). Advances made to reduce the 

fuel consumption, noise, and emissions of these aircraft could provide a considerable positive 

contribution to the goal of minimizing the future environmental impact of aviation. 

2.0 Study Objectives and Approach 

This is a preliminary study with the objective of assessing the potential benefits and penalties of open 

rotor propulsion concepts as applied to an advanced 737/A320 replacement aircraft. This initial study is 

necessarily limited in detail and scope due to a lack of analysis tools to address all aspects of open rotor 

propulsion design, performance, and integration. Models used are approximate and the results should be 

viewed as preliminary while enhancements to the methods and data used to feed the processes continue. 

The general approach taken for this study is to develop analytical models of advanced engines, combine 

them with an advanced technology airframe, design the overall system to meet mission requirements and 

constraints, and assess the resulting noise, fuel consumption, and emission characteristics. 

3.0 Modeling and Analysis Methodology 

3.1 Propulsion Modeling 

Propulsion system modeling was performed using the Numerical Propulsion System Simulation 

(NPSS) code (refs. 11-13) for thermodynamic cycle analysis and performance and the Weight Analysis of 

Turbine Engines (WATE) code (refs. 14-16) for aeromechanical design and weight estimates. Estimates 

for emission of nitrogen oxides (NOX) were obtained from an empirical correlation representing an 

advanced, low-NOX combustor. Reference 17 provides more details on this empirical NOX correlation, 

which was developed by NASA combustor technologists during the latter stages of NASA’s UltraEfficient Engine Technology Program. Developing the capability to design and analyze open rotor 

engines with NPSS was a significant task in itself. A new counter-rotating propeller performance element 

was added to NPSS to enable proper modeling. To model rotor performance, performance maps were 

generated from experimental data obtained for a General Electric ”Gen-1” rotor design at high speeds 

(NASA Glenn 8x6-ft Supersonic Wind Tunnel) and low speeds (NASA Glenn 9x15-ft Low-Speed Wind 

Tunnel). Although the capability to model both geared and direct drive open rotor engines was established 

in NPSS, a geared open rotor architecture has been assumed for this assessment based on an internal 

concept evaluation and input from General Electric. References 18 and 19 provide more details on the 

open rotor modeling approaches that have been recently developed. Engine technology assumptions for 

the current study were based on those used in the NASA advanced turbofan engine design studies 

described in references 17 and 20. Details on the analysis methodology for the geared turbofan engine 

model used as a comparison case in this study are also provided in references 17 and 20. 
4 

3.2 Airframe Modeling 

Previous open rotor studies have identified a number of possible airframe integration options. 

Reference 21 contains an excellent discussion of the various propulsion system and integration options, 

and the benefits and issues associated with each combination. The primary focus toward the end of the 

open rotor work of the 1980s was the rear-mounted pusher configuration (see Figure 2). As detailed in 

reference 21, this integration approach is favored from the aspect of passenger accommodation (lower 

interior noise) and wing aerodynamics. Although there are a number of issues that would need to be 

addressed to make this configuration successful, such as foreign object damage susceptibility, stability 

and control impacts, and pylon wake-rotor interaction, this layout seemed to be the preferred approach 

during the late 1980s. Because NASA’s current focus in open rotor propulsion is the counter-rotating, 

pusher concept and a rear-mounted installation is the most likely configuration for this propulsion system, 

a rear fuselage-mounted engine layout was selected for this initial study. 

Figure 2. Rear-mounted open rotor pusher configuration, shown on Boeing 727 test vehicle. 

Development of a suitable airframe analytical model began with modeling of a rear-engine, singleaisle MD-90-30 type configuration in the NASA aircraft sizing and synthesis code FLOPS (Flight 

Optimization System) (ref. 22). This model was developed using publicly-available data on weight and 

performance of the MD-90-30 (refs. 23, 24) and a NASA-developed engine model representative of the 

International Aero Engines V2525-D5 engine. The wing and fuselage structural weight estimates of 

FLOPS were replaced with estimates from PDCYL (ref. 25), which offers a less empirical, more 

analytical weight estimation methodology. Using geometry, weight (maximum gross weight and 

maximum landing weight), and mission data from reference 23, and wing weight and fuselage weight 

predictions from PDCYL, FLOPS predicts an operating empty weight (OEW) of 91,470 lb compared to 

the manufacturer’s reported value of 88,175 lb. Although this represents only a 3.5% difference between 

predicted and actual weight, the FLOPS model was calibrated to the actual value by applying an overall 

reduction factor on empty weight. Calibration could not be performed for each individual component 

since actual weight data were not available at that level. A point on the payload-range diagram of 

reference 23 was used to calibrate the range performance of the FLOPS model. Assuming 31,000 lb of 

payload (155 passengers) and a takeoff gross weight of 156,000 lb, reported range capability for the MD90-30 is approximately 2040 nm. The initial FLOPS-predicted range capability was significantly higher. 

There are a number of possible explanations for the range discrepancy, including using a different mission 

profile (cruise Mach, altitude, and reserve mission assumptions), differences in predicted and actual 

aerodynamic performance, and differences between the NASA engine model and actual engine fuel 

consumption characteristics. Without additional data in one of these areas, it is impossible to determine 
5 

the underlying cause. Because the FLOPS-predicted cruise aerodynamic efficiency did not seem 

optimistic and the cruise thrust specific fuel consumption (TSFC) seemed low compared to other engines 

of this class, range performance of the model was calibrated to the published value by adjusting the fuel 

flow rates of the V2525-D5 engine model. In reality, the range discrepancy is likely due to a combination 

of factors. 

A series of modifications was necessary to progress from a MD-90-30-like FLOPS model to an 

appropriate advanced technology airframe model suitable for the open rotor engines. These steps are 

summarized in Figure 3 and described in more detail below. 

MD-90-30 Like Model

Improve Wing 

Aerodynamics

Lengthen Fuselage for 

162 Passengers

CSAT-re

Resize for 3250 nm 

Mission

MD92V Model MD92 Model

Open Rotor Airframe 

Impacts Advanced Airframe 

Technology Assumptions

ASAT-re ASAT-or

Figure 3. Process used to model advanced technology airframe suitable for open rotor propulsion. 

Previous NASA studies of propulsion options for 737/A320 replacement designs (refs.17, 20) have 

used a Boeing 737-800-based airframe model and design mission. Although the MD-90-30-based model 

described above differs from the 737-800-based model in a number of respects, it was desired to at least 

have a consistent technology level and design mission for the baselines used in the two studies. Of critical 

importance in developing a consistent technology baseline is the level of airfoil design technology for the 

MD-90-30 wing. The MD-90-30 uses the same wing design as the older MD-80 series. This 1980s 

technology is not as advanced as the newer 737-800 wing (1998 entry-into-service). Although the exact 

difference in aerodynamic design technology between the 737-800 and MD-90-30 is not known in the 

public domain, there is some indication of the potential gap in aerodynamic efficiency provided by 

comparing the current 737 design to older models. One source indicates maximum M*(L/D) (Mach 

number times lift-to-drag ratio) of the 737-700 is 16% higher than the equivalent size, older generation 

737-300 design (ref. 26). In order to advance the MD-90-30-like airframe model to late 1990s technology 
6 

levels, FLOPS aerodynamic factors were adjusted to provide a similar improvement in aerodynamic 

efficiency. The takeoff and landing aerodynamic performance was also modified. A model of the 737-800 

low-speed performance was used as a rough approximation for the notional, updated version of the MD90-30. Changes were made to the design mission to make it consistent with the prior studies. The study 

design mission (162 passengers, mixed class, 3250 nm) was selected to reflect performance enhancements 

projected for future advanced aircraft in this vehicle class. Although based on desired future capabilities, 

this design mission was also applied to the older technology baseline model to enable a consistent 

comparison with the advanced study vehicles. Since the MD-90-30 fuselage only accommodates 158 

passengers in mixed class seating, the fuselage length was slightly increased to accommodate an 

additional row of seats. Wing area was increased to accommodate the higher gross weight needed to meet 

the design mission. The resulting airframe, referred to as CSAT-re (Current technology Single-Aisle 

Transport-rear engine), represents a model of the MD-90-30 updated to late 1990s entry-into-service 

technology levels and re-sized to perform the study design mission. 

The Advanced Single-Aisle Transport-rear engine (ASAT-re) airframe model is a derivative of the 

CSAT-re model intended to be representative of a potential advanced technology single-aisle replacement 

aircraft. The primary airframe technology advancement assumed is extensive use of composite materials 

for the airframe structure. The composite construction is assumed to result in a 15% reduction in weight 

of the wing, fuselage, and empennage compared to the metal construction used for the CSAT-re airframe. 

This weight reduction represents an assumed benefit for future composite structures and is not necessarily 

a reflection of the benefit composite materials have provided to date. Other minor technology 

improvements include an increase in hydraulic pressure to 5000 psi, and a 1% reduction in drag 

associated with a variable camber trailing edge and general drag clean-up. The basic geometry parameters 

for the ASAT-re were not changed from the CSAT-re model. 

The MD-90-30-based ASAT-re airframe layout is generally compatible with an open rotor propulsion 

system, and, in fact, open rotor propulsion was considered during the design of the MD-90 family. 

Reference 27 describes the design of both turbofan and open rotor configurations for the MD-90 family 

and details the modifications to the turbofan version that are necessary to accommodate an open rotor 

engine. To approximate the impact of those modifications, the MD-92 configurations described in 

reference 27 were examined and modeled in FLOPS. Reference 27 discusses several differences between 

the MD-92 (open rotor) and MD-92V (turbofan) airframe designs necessitated by the integration of the 

open rotor system. The fuselage weight is higher due to additional acoustic treatment for reduced interior 

noise, sonic fatigue treatment of the aft fuselage structure, and structural treatments to reduce vibrations. 

Hydraulic system weight is increased by an additional hydraulic system dedicated to the tail control 

surfaces. A center-of-gravity management system is needed to address stability and control issues 

introduced by the heavier aft engine weight and higher engine thrust line. The system-level impact of 

these differences, plus the higher weight of the open rotor engine, was a 4200 lb increase in empty 

weight. Since detailed weight increments were not provided in reference 27, this overall weight increase 

was allocated in an approximate manner to each of the above areas to develop a FLOPS model of the 

MD-92 from a MD-92V FLOPS model. The adjustments necessary to convert from the MD-92V 

(turbofan) to the MD-92 (open rotor) configuration were then applied to the ASAT-re model described 

above to develop the ASAT-or (Advanced Single-Aisle Transport-open rotor) model for use in the open 

rotor aircraft analysis. The nominal ASAT-or configuration is shown in Figure 4. 
7 

Figure 4. Basic Advanced Single-Aisle Transport configuration with rear-mounted open rotor engines (ASAT-or). 

3.3 Noise Analysis 

Prediction of FAR Part 36 certification noise (ref. 8) for open rotor propulsion is of critical importance 

in understanding the trades between lower fuel consumption and higher noise. Open rotor noise prediction 

is also an area with significant challenges. A number of counter-rotating open rotor source noise 

prediction methods were developed during NASA’s previous studies of open rotors in the 1970s and 

1980s. Perhaps the most promising of these methods are the semi-empirical codes described in references 

28 and 29, which have recently been revived and validated on modern computing platforms. However, 

these methods predict only rotor-rotor discrete interaction tones and they are incapable of evaluating rotor 

broadband noise. They are also not capable of evaluating the noise of rotors having advanced, low-noise 

design features. Open rotor blades are now designed using modern, 3D aerodynamic analytical tools and 

advanced rotors will employ low-noise features, some of which have been made public and others that are 

still company-proprietary. Known rotor noise mitigation strategies include aeroacoustic blade shaping, 

blade pitch angle and rotational speed optimization, increased blade counts, low disk loading, rotor-rotor 

and pylon-rotor spacing optimization, aft rotor clipping, and pylon wake reduction (ref. 30). Because of 

the shortcomings of current rotor noise modeling capabilities, a system-level noise prediction method has 

been developed which uses noise data measured from scale model open rotor test articles. 

3.3.1 Open Rotor Source Noise Modeling 

Tests of various open rotor designs have taken place in NASA Glenn’s experimental facilities as part 

of an ongoing NASA-General Electric collaborative partnership (ref. 31). Low-speed aerodynamic and 

acoustic tests were conducted in the Glenn 9x15-ft Low Speed Wind Tunnel, and high-speed 

aerodynamic testing was conducted in the Glenn 8x6-ft tunnel. The open rotor acoustic data used for this 

assessment of certification noise was collected from one of General Electric’s advanced “Gen-1” rotor 

blade sets. 

The experimental setup is illustrated in Figure 5, showing a schematic of the open rotor rig and the 

linear microphone array. The microphone traverse array is located 60 inches from the centerline of the 

rotor model. Each tunnel acoustic “reading” consists of microphone-corrected, narrowband spectral 

density levels taken at 18 sideline microphone locations with frequencies ranging from zero to 100 kHz at 

frequency intervals of 12.2 Hz. Data were collected with the model operating at various shaft speeds and 

blade pitch angles. Both rotors were always operated at identical shaft speeds. Tests were also conducted 
8 

to estimate the facility’s tare background noise levels with no power supplied to the open rotor test article 

and with no blades attached. 

Figure 5. Open rotor test article and linear microphone array in the NASA Glenn 9x15-ft Low Speed Wind Tunnel. 

(Diagram courtesy of Edmane Envia, NASA Glenn Research Center)

The microphone-corrected spectral density levels were adjusted by the facility’s data acquisition 

software to lossless, freefield conditions using an atmospheric attenuation model. The spectral density 

levels were further corrected for spherical spreading from the microphone’s constant sideline distance to a 

1-foot constant radius arc from a point originating on the rear rotor center of rotation. These data must be 

manipulated in several additional ways in order to properly predict certification noise for a full-scale 

aircraft. The spectra were modified further by applying the following corrections: 

1. Initial preparation 

The spectral density levels of the tare test were subtracted from each noise spectrum of interest in the 

powered tests over all frequencies. Next, the spectral densities were converted to sound pressure levels 

with a reference pressure of 20μPa. The geometric sideline microphone array angles (G) range from 18° 

to 140°, where the zero reference of G is directly upstream. Emission yaw (polar) angles (E) were 

computed from the G angles using the relation E = G - sin-1[MTunnel sin G], where the tunnel Mach 

number, MTunnel, in the experiments is 0.20. A small amount of data were collected for MTunnel = 0.18 and 

0.22, but they were not used in this assessment. 

2. Conversion to static conditions 

The sound pressure levels were corrected for convection effects to static conditions by adding an 

amplitude correction of 10 log10[ 1 - MTunnel cos E ]SME, where the source motion exponent (SME) was 

taken to be 2, for dipole emission. No Doppler adjustments were made to the frequency scale (i.e., fStatic = 

fTunnel), since the microphones were “traveling with” the source at the speed of the tunnel Mach number.

3. Conversion to flight conditions 

The sound pressure levels were corrected to realistic takeoff and approach flight conditions by 

subtracting an amplitude convection correction of 10 log10[ 1 - MFlight cos E ]SME. The flight Mach 

number, MFlight, is supplied via airplane trajectory calculations for approach and departure, and is

discussed in later sections. The source motion exponent was again taken to be 2. A Doppler frequency 

shift (fFlight = fStatic / [ 1 - MFlight cos E ]) was applied to the frequency scale. In addition, an amplitude 

adjustment of 10 log10[ MFlight / MTunnel ] was made to account for observed increases in source levels with 

freestream Mach number. (This correction is based on experiments conducted by General Electric during 

the Advanced Turboprop Project.) Further, a source strength amplitude adjustment of 

10 log10[(Flight/Tunnel)

2

(cFlight/cTunnel)

4

] was made to correct the spectra from tunnel conditions to flight 
9 

conditions, where Tunnel and cTunnel are the density and the speed of sound measured in the wind tunnel, 

respectively, and Flight and cFlight are the density and the speed of sound at the simulated flight conditions. 

Standard acoustic day (ISA+18°F) values at altitudes predicted by the trajectory analysis were used for 

the flight conditions.

4. Conversion to full scale 

The sound pressure levels were then corrected from model scale to full scale. Frequencies were shifted 

to lower values by dividing by the linear scale factor, DFull-Scale/DRig, and amplitudes were adjusted by 

adding a correction of 20 log10[ DFull-Scale/DRig ]. The scale model test article front rotor diameter, DRig, was 

2.4095 ft. DFull-Scale depends on the characteristics of the study engine (Section 4.1) and any engine thrust 

scaling performed to match the engine and airframe (Section 4.2). The final, sized front rotor diameter for 

the airplane in this study is 13.67 ft resulting in a linear scale factor (DFull-Scale/DRig) of 5.673.

5. Conversion to 1/3rd octave band spectra 

The frequency basis for FAA and ICAO aircraft noise certification is the traditional 1/3rd octave band 

center frequency spectrum, represented by 24 discrete sound pressure levels ranging from 50 Hz to 

10,000 Hz. The acoustic energy contained in the intervals on either side of each center frequency (i.e., 

from 2-1/6 times the center frequency to 21/6 times the center frequency) are summed to form each of the 24 

standard 1/3rd octave band sound pressure levels. The band filter specified by Part 36 noise regulations is 

defined by the International Electrotechnical Commission (ref. 32). 

Several of the processing steps discussed above are illustrated in Figure 6. For illustration purposes, 

spectra from a non-proprietary – but representative – rotor set were used. The sample spectrum was taken 

from a microphone reading in the forward quadrant. A one-foot-arc, lossless, narrowband, sound pressure 

level spectrum with facility noise removed is shown by the black line in Figure 6a. This spectrum has an 

emission yaw angle of 24° from the inlet, the tunnel Mach number was 0.20, and the open rotor test 

article was operating at a corrected shaft speed of 6300 rpm. Although noise measurements are available 

up to 100 kHz, only lower-frequency data are plotted. When flight condition scaling is applied (scaling to 

a flight Mach of 0.26 at 1000 ft is shown), the spectrum shifts to higher amplitudes and frequencies, as

represented by the red line in the figure. Opposite shifting behavior occurs for spectra in the aft quadrant. 

Scaling from experimental tunnel conditions to appropriate, simulated flight conditions can result in 

changes of several decibels in the fore or aft quadrants and is thus an important processing step. When the 

scale factors are applied to convert from model scale to full scale, the spectrum shifts to much higher 

amplitudes and lower frequencies represented by the blue lines in Figure 6a and Figure 6b. At one-fifth 

scale, the inaudible sound pressure levels measured at frequencies of about 50 kHz are shifted to the 

audible, regulated range used to estimate noise certification levels. 

The 1/3rd octave band sound pressure levels are plotted in Figure 6b. The fundamental blade passage 

frequency tones for the front and aft rotors are labeled “1f” and “1a” in the figure, respectively. Since the 

front rotor has 12 blades and the aft rotor has only 10 blades in this particular design, the fundamental 

tone of the front rotor is always at a higher frequency than the fundamental tone of the aft rotor. Unequal 

blade counts distribute the tones over a wider frequency range and help reduce noise. Several higher-order 

rotor-rotor discrete interaction tones are also noted in the figure. Seldom do any of the interaction tones 

above the 2f+2a tone contribute significantly to the overall sound pressure level. The reasoning behind 

the frequency “coarsening” from the narrowband to the 1/3rd octave band is that the resulting spectra 

become much easier to manipulate and store. However, many of the rotor-rotor interaction tones lie 

within a single frequency interval, such that a single 1/3rd octave band sound pressure level may 

incorporate two (or more) tones. This is particularly true at higher frequencies since the scale is 

logarithmic. 
10

7 0

8 0

9 0

100

110

120

130

140

150

1 0 100 1000 10000

SPL, dB

Frequency, Hz

Processed and Scaled Narrowband

Processed and Scaled 1/3rd Octave Band

1 a

1 f

1f+1a

1f+2a

2f+2a 5 dB

7 0

8 0

9 0

100

110

120

130

140

150

0 1000 2000 3000 4000 5000

SPL, dB

Frequency, Hz

Data with Facility Noise Removed

Brought to Simulated Flight Conditions

Scaled to Full Size

 a) Conversion to full scale. b) Conversion to 1/3rd octave band. 

Figure 6. Sample open rotor narrowband spectrum conversions. 

(MTunnel =0.2, E =24°, corrected shaft speed=6300 rpm) 

3.3.2 System-Level Noise Predictions 

The Aircraft Noise Prediction Program (ANOPP) Release Level 30 (refs. 33, 34) is a systems-level 

code used in this study to compute certification noise for the open rotor airplane. The center frequency 

basis for ANOPP is the 1/3rd octave band. The scaled 1/3rd octave band open rotor spectra described 

above were fed into ANOPP via its Acoustics Data Module (ACD). ACD is an ANOPP utility that allows 

user-supplied spectra to be fed into a certification simulation in lieu of using ANOPP’s own, built-in, 

source noise prediction modules (ref. 35). Core, jet, and airframe noise sources were modeled using the 

ANOPP methods described in references 36, 37, and 38, respectively. Using an assumption of acoustic 

superposition, the freefield, lossless spectra for all of the noise sources were analytically summed in the 

vicinity of the aircraft. Real noise sources are, of course, complex, distributed signals that are affected by 

other acoustic sources, aircraft external surfaces, and the environment. No attempt was made to adjust the 

component spectra for acoustic near-field phenomena such as source interactions, reflections, refraction, 

diffraction, or other effects. 

An airplane model with appropriate low-speed aerodynamics, weight, and propulsion performance 

data is necessary to model the takeoff and approach trajectories needed for aircraft certification 

predictions. The airplane was modeled using the FLOPS computer program as described in Section 3.2. 

FLOPS computes detailed, low-speed takeoff and landing profiles using a built-in, time-stepping 

trajectory analysis module. A takeoff and landing analysis was performed with FLOPS for a sea level 

field, with zero gradient, at standard acoustic day (ISA+18°F) conditions using the vehicle’s low-speed 

aerodynamics and the open rotor thrust data. Compliance with the airworthiness requirements described 

in Parts 36 and 25 of the Federal Aviation Regulations (references 8 and 39, respectively) was observed. 

For takeoff, maximum takeoff gross weight was assumed and the flaps were set at detent 5. The takeoff 

procedures were as follows: 

1. Brake release at maximum Sea Level, Static (SLS) rated thrust 

2. Ground run 
11

3. Rotate such that liftoff velocity is at least 110% of the minimum unstick velocity with all engines 

operating, or 105% of minimum unstick velocity with one engine inoperative 

4. Climb to the 35 ft obstacle at maximum thrust 

5. Raise gear 

6. Climb at constant calibrated airspeed at maximum thrust 

7. From 16,000 ft to 17,000 ft from brake release, reduce thrust to the minimum allowable cutback 

value (such that climb gradient is zero with one engine inoperative, or 4% with all engines 

operating) 

8. Climb at constant calibrated airspeed at cutback thrust 

9. At 30,000 ft from brake release, increase thrust to “power-up” level 

10. Climb at constant calibrated airspeed at power-up thrust 

For approach, a 3° glide slope was followed, the maximum landing weight was assumed, and the flaps 

were set at detent 40 with leading edge slats and landing gear deployed. 

Vector geometry analyses for the airplane relative to the three certification microphone measurement 

locations – shown in Figure 7 – were performed within ANOPP as functions of source time. The spectra 

were then propagated to the three certification observers on the ground in accordance with the 

specifications for certification measurements. 

2000 m

(6562 ft)

Flyover (with cutback)

Reference

Lateral

Reference

Approach

Reference

6500 m

(21 325 ft)

450 m

(1476 ft)

Figure 7. Noise observer arrangement relative to takeoff and landing aircraft trajectories. 

Noise propagation effects accounted for included spherical spreading, Doppler shift and convective 

amplification, atmospheric absorption (ref. 40) and variable impedance effects within a layered 

atmosphere, ground reflections (ref. 41) based on data for grass-covered ground (ref. 42), and lateral 

attenuation (ref. 43). To account for incoherently-radiating twin engines, 3dB were added to the 

propulsion noise sources. More complex propagation phenomena such as scattering, weather effects, and 

terrain were not modeled. 

From the propagated spectra, ANOPP computes several noise metrics of interest as functions of 

observer time. The Effective Perceived Noise Level (EPNL) certification noise metric is computed from 

the noise-time history at each observer as prescribed in reference 8. 

3.3.3 Special Considerations for Open Rotor Aircraft 

In a propeller-driven aircraft noise certification test (or a simulation of one), the tone-rich acoustic 

spectra observed by the microphones fluctuate significantly when tones move from one frequency interval 
12

into another as the aircraft-observer geometry changes in flight, and convective amplification and Doppler 

effects come into play. This leads to irregular tone-corrected, perceived noise level (PNLT) vs. time 

histories as the airplane traverses a path relative to its certification observers. These irregularities are 

enhanced when ground reflections are present – causing ground nulls and pseudotones – particularly at 

frequencies under about 800 Hz, where ground effects are most efficient and where many propeller tones 

exist (as can be seen in Figure 6b). This effect is observed for current large turboprops in the transport 

category and may be observed one day for open rotor airplanes. 

A special computational procedure unique to transport-category propeller aircraft was developed to 

handle the tone correction penalties from spectra that are rich in both real rotor tones and fictitious 

pseudotones. If tones are propagated to an observer, regulations (ref. 44) permit tone correction penalties 

resulting from them to be ignored provided the tones can be shown to be unrelated to engine source noise, 

since they may be artifacts of ground reflections. When PNLTs are analytically computed in this study, 

two ANOPP runs are made. In the first run, the sources are analytically flown past the observers, the 

ground reflection calculations are turned off, and the tone correction penalties at half-second intervals are 

saved. In the second run, the sources are again flown past the observer, the ground reflection calculations 

are turned on, and the PNL metric is computed at half-second intervals. The tone correction penalties 

from the first run are added to the PNLs from the second run, and the result is integrated to form the 

certification metric, the EPNL. Thus, this process is effective at excluding pseudotones from the tone 

correction calculations. This procedure is believed to be consistent with the regulations found in Ref. 44, 

but it has not been verified with regulatory authorities. 

Another issue related to the tone correction penalty is the location of the lateral observer. Propeller 

tones were reason that the lateral observer location for large, transport-category, propeller-driven 

commuter transports was changed in 2002 from the 450 m (1476 ft) lateral sideline to a flyover point 

directly under the flight path where the airplane has reached an altitude of 650 m (2133 ft). (See Section 

B36.3(a)(2) of Part 36 (ref. 8) beginning in 2003.) The lateral observer relocation was permitted because 

of the difficulty inherent in determining the location of peak noise on the lateral sideline for turboprop 

aircraft due to the large number of ground nulls and pseudotones caused by ground reflections. With the 

observer located on the extended runway centerline, ground effects are less efficient because the airplane 

is flying directly overhead and the angles that the source subtends with the horizon are greater. In this 

study, however, the lateral observer location for the NASA notional open rotor airplane is assumed to 

remain on the 450 m lateral sideline (as required for jet aircraft in Section B36.3(a)(1) of Part 36, ref. 8). 

This is because an open rotor vehicle will likely be larger than commuter turboprops, and it will naturally 

be compared to (and compete with) contemporary, large, single-aisle jet aircraft. At this writing, however, 

placement of the lateral observer remains an open issue for large, open rotor transports. With lateral 

attenuation modeling included, the peak EPNL along the lateral sideline occurs across from the point 

where the airplane reaches an altitude of about 900 ft above field elevation. 

The approach and flyover observers are located on the extended runway centerline as prescribed by 

reference 8, with the approach observer located 2000 m (6562 ft) from the runway threshold (or, 

assuming a 50 ft landing obstacle and a 3 degree glide slope, 2291 m (7518 ft) from the touchdown 

point), and the flyover observer located 6500 m (21,325 ft) from brake release. See Figure 7 for a sketch 

of these locations. 

3.3.4 Establishing Reference System Noise Levels 

Seventy-two acoustic readings collected from the NASA Glenn 9x15-ft Low Speed Wind Tunnel were 

used in establishing reference system noise levels. These readings, made during May 2010, had an 
13

isolated nacelle (i.e., no pylon or simulated fuselage was used) and a muffler system to effectively 

suppress the noise of the drive rig. All 72 readings were made at a Mach number of 0.20. Data were 

collected at a variety of blade pitch angles and shaft speeds. Front rotor advance ratios (varied by 

changing the rotor shaft speeds) ranged from 0.82 to 1.48. 

The acoustic readings were scaled and analytically flown past the observers as described in earlier 

sections and EPNLs were computed for each of them. An example of this procedure is shown in Figure 8

for the flyover observer. In this case, the open rotor is operating at a corrected tip speed of 655 ft/s with a 

front blade pitch angle of 36.8°. The linear scale factor, flight Mach number, and altitude used to scale the 

spectra are 5.673, 0.269 and 2050 ft, respectively. The PNLTs are plotted against the observer time from 

brake release at half-second intervals (left), and against the polar emission angle (right). The EPNL 

(81.0 EPNdB in this case) is computed by integrating the PNLT vs. time curve using only the points 

within 10 PNdB of the maximum. The jet, core, and airframe noise sources also contribute to the overall 

levels. Jet noise, however, is virtually negligible in all cases due to the low core jet velocities exiting the 

power turbine. Airframe and core sources contribute only at approach, when rotor noise levels are 

relatively low. 

In the tunnel, microphone readings were taken from polar geometric angles ranging from 18° to 140°. 

With the tunnel operating at Mach 0.20, the polar emission angles range from 14° to just 133°. 

Unfortunately, when the sources fly past an observer, noise remains within 10 PNdB of the maximum 

until the source recedes beyond 150°. There is a region in the “10 PNdB-down” area where the emission 

angles are greater than the aft most microphone reading of 133°. In this region, the aft most microphone 

spectrum is simply repeated as the aircraft recedes until spherical spreading diminishes its level. 

72

74

76

78

80

82

84

85 90 95 100 105 110

PNLT, PNdB

Observer Time from Brake Release, s

40 60 80 100 120 140 160

Emission Angle, deg

Figure 8. Example flyover observer PNLT noise histories for a case with rotor corrected tip speed of 655 ft/s. 

The set of 72 EPNLs correlate very well with several independent variables such as rotor thrust, tip 

speed, and blade pitch angle. To obtain the reference system EPNL for each of the three certification 

observers, the characteristics of the airplane as it flies past each observer must be matched against the 

EPNL dataset. In other words, the EPNL for each observer is selected based on the thrust demand and 

flight conditions of the airplane. The thrust, altitude, airspeed, and angle of attack are known via the 

takeoff and approach trajectory calculations described in Section 3.3.2 and shown later in Section 4.3.1. 

This thrust-noise matching procedure requires that the thrust levels measured by the wind tunnel open 

rotor propulsion rig be scaled to the thrust that an actual full-size open rotor propulsion system would 

produce in flight. Thrust scaling is a simple proposition for a propeller that maintains geometric and 

aerodynamic similarity throughout the takeoff and landing flight regime. However, in practice, variable-
14

pitch propellers operate in flight conditions that are decidedly not similar to the tests conducted in the 

tunnel. For practical engine operational and performance reasons, the rotor tangential tip speed (and shaft 

speed) is held constant. Rotor blade pitch angle is varied as flight conditions change to provide the thrust 

required. As the airspeed increases from zero at brake release to climbout velocities of well over 150 kts, 

the advance ratio increases as well. Blade pitch angles must increase with airspeed to maintain peak 

efficiency. With a constant-speed propeller, scaling thrusts from the wind tunnel Mach number (0.20) to 

the flight Mach number (often in excess of 0.26) is required. 

No theoretical scaling approach is perfect in non-similitude cases. Ideally, if there were sufficient wind 

tunnel data collected at a variety of low-speed flight Mach numbers, a simple empirical thrust correction 

factor model could be developed. However, nearly all of the open rotor data were collected at a wind 

tunnel Mach number of 0.20. Therefore, instead, a simple thrust correction multiplier is required that 

performs reasonably well for minor excursions in flight conditions in the takeoff regime. This correction 

is discussed in greater detail in the Appendix. The correction adjusts the rotor thrusts measured in the 

tunnel to the appropriate flight conditions by applying multipliers of the density ratio, diameter ratio, and 

a factor that accounts for the inlet ram drag and core nozzle thrust of the engine. The results of the thrustnoise matching procedure will be presented in Section 4.3.2. 

3.3.5 System Corrections to Reference Noise Levels 

Several important system-level adjustments must be made to each of the reference EPNLs. Since the 

reference EPNLs represent the noise signature of an isolated open rotor nacelle in a uniform flowfield 

with steady loading, corrections are required to account for the effect of the rotor inflow angle, the 

presence of an upstream engine pylon, and pylon wake mitigation strategies envisioned by NASA and 

General Electric. Other system-level effects not accounted for in the current analysis include fuselage and 

tail shielding effects on the lateral observer, other fuselage effects, T-tail reflections, and flowfield effects 

due to rotor rotation direction. 

The open rotor wind tunnel model was tested at rotor inflow angles of 0°, 3°, and 8° by angling the rig 

into the freestream flow. Rotor noise was found to be a strong function of rotor inflow angle due to the 

unsteady aerodynamic loading effects on the rotor blades. To use the wind tunnel data collected at 

nonzero angles to account for inflow effects, the rotor inflow angle, Inflow, must first be related to the 

angle of attack of the airplane,  (known from the trajectory analysis). A vortex-lattice aerodynamic 

analysis was performed to estimate the angularity of the flow at the rotor location. For this linear analysis, 

Inflow may be expressed as: 

Inflow Cant 0

   /1 dd 

where Cant is the propulsion system mounting cant angle (relative to the aircraft zero angle of attack 

line),  is the downwash angle at the rotor when  = 0, and d/d is the variation of downwash with 

airplane angle of attack. The parameters  and d/d are functions of the wing configuration (i.e., they 

vary with the amount of flap and slat extension). 

To set the cant angle, the vehicle was first analyzed in a clean configuration at cruise flight conditions 

and the downwash angle into the rotor face was computed. The cant angle, Cant, was then set to 2° (noseup) such that the rotor inflow angle is approximately zero at the cruise flight condition. The vehicle then 

was analyzed in departure (partial flaps deployed) and approach (full flaps and slats deployed) 

configurations. On departure,  and d/d were found to be 2.34° and 0.336, respectively. On approach, 

 and d/d were found to be 5.19° and 0.349, respectively. Using the  computed during the aircraft 
15

trajectory analysis, the approximate rotor inflow angle was computed at each observer and an appropriate 

inflow angle noise penalty assigned to each EPNL. 

The increase in rotor noise due to the pylon wake impinging on the front rotor was estimated by an 

examination of data from testing of General Electric’s historical baseline rotor design in the presence of 

an upstream engine pylon. Only eight readings were taken with the pylon in place. Four readings were 

made at zero inflow angle, and two readings were made at inflow angles of 3° and 8°. With this limited 

amount of data available, it was observed that the greatest pylon noise impact occurred at zero inflow 

angle and at low thrust levels. There was virtually no distinguishable impact at the 8° inflow angle. Some 

of the pylon noise impact can be reduced by applying pylon wake mitigation technologies under 

consideration by NASA and General Electric. To account for the presence of a pylon with successful 

implementation of pylon wake mitigation strategies, penalties of 0.6, 0.3, and 0.6 EPNdB were assigned 

to the approach, lateral, and flyover EPNLs, respectively. 

4.0 Analysis Results 

4.1 Engine Design 

General characteristics of the NASA advanced open rotor engine are shown in Table 1, additional 

details are provided in reference 19. Also shown for comparison are the characteristics of one of NASA’s 

advanced geared turbofans from reference 20 and NASA’s model of the V2525-D5 turbofan, which 

powers the MD-90-30 aircraft used as the starting point for the open rotor aircraft concept. Four different 

geared turbofan designs were developed in reference 20. The geared turbofan selected for comparison 

here is the fan pressure ratio (FPR) 1.5 engine, which resulted in the lowest mission fuel consumption in 

that study. The open rotor engine was designed for 5,000 lb of thrust at top-of-climb (TOC) conditions 

(M=0.78, 35,000 ft) and 19,000 lb of thrust at hot day, rolling takeoff conditions (M=0.25, SL). The open 

rotor engine was initially designed for 17,800 lb of thrust at hot day, rolling takeoff conditions. However, 

the operating temperatures at cruise for the resulting engine were too high. As a result, the final open rotor 

engine is “oversized” for takeoff. This excess thrust can be used for improved takeoff performance, or the 

engine could be “de-rated” for takeoff which would improve engine life. The geared turbofan design 

conditions are slightly different: 5,000 lb at M=0.8, 35,000 ft and 17,500 lb at the hot day, rolling takeoff 

condition. The TOC thrust and TSFC values in Table 1 are given for the M=0.78 open rotor TOC 

condition for a consistent comparison. The advanced geared turbofan and open rotor engines are shown 

graphically in Figure 9. 

The expected efficiency benefits of the advanced open rotor architecture are clearly evident in Table 1. 

At the M=0.78, 35,000 ft top-of-climb point, the open rotor engine TSFC of 0.428 lb/(lb·h) is 

approximately 13% lower than for the FPR=1.5 geared turbofan (0.494 lb/(lb·h)) and 30% lower than for 

NASA’s V2525-like model (0.610 lb/(lb·h)). The reduction in TSFC is even greater for the SLS and 

rolling takeoff conditions. At SLS conditions, the open rotor TSFC of 0.158 lb/(lb·h) is 39% lower than 

for the advanced geared turbofan and 54% lower than for the V2525-like engine. Similar reductions in 

TSFC are obtained at rolling takeoff conditions. These improvements in TSFC at SLS and rolling takeoff 

should lead to much lower fuel consumption and emissions in the airport terminal area. Landing-Takeoff 

(LTO) NOX characteristics are presented in Table 1 in terms of the regulated parameter, Dp/Foo, where Dp

is the grams of NOX emitted over a standard LTO cycle (by a single, uninstalled engine) and Foo is the 

rated output at SLS conditions in kilonewtons. This parameter is defined by ICAO and used in FAR Part 

34 for engine certification (ref. 45). Dp/Foo for the open rotor engine is 12.15 g/kN, 78% less than for the 

V2525-D5. The reduction in LTO NOX emissions compared to the V2525 engine comes from a 

combination of lower NOX emission indices (EIs) due primarily to an advanced low NOX combustor and 
16

lower fuel flow (higher efficiency) at the ICAO-defined LTO operating points. The advanced geared 

turbofan engine has the same advanced combustor as the open rotor, and similar NOX EIs. However, the 

significantly lower fuel flow of the open rotor engine leads to a 45% reduction in Dp/Foo compared to the 

advanced geared turbofan. The estimated NOX Dp/Foo of 12.15 g/kN for the open rotor engine is 79% 

below the current ICAO regulatory limit, established at the sixth meeting of the Committee on Aviation 

Environmental Protection. 

The reduction in TSFC and LTO NOX does not come without penalty. Despite the use of advanced 

materials in the open rotor engine, the thrust-to-weight ratio (T/W) of the open rotor engine is actually 

lower than for the older technology V2525-D5 turbofan. The T/W of the open rotor engine at top-ofclimb is 0.54, 25% lower than for the V2525-like model and 28% lower than for the advanced geared 

turbofan. At the aircraft system level, the benefit of the large TSFC reduction is offset somewhat by the 

lower engine thrust-to-weight, as well as the additional airframe weight associated with accommodating 

open rotor propulsion. 

Table 1. General Engine Characteristics 

V2525 Turbofan 

Model

Advanced Geared 

Turbofan (FPR=1.5) Geared Open Rotor

Overall Length, ft 18.3 15.8 23.2

Nacelle Diameter, ft 8.0 7.6 5.6

Rotor Diameter, ft (front / rear) N/A N/A 13.8 / 13.1

Engine+Nacelle Wt., lb 7900 6626 9218

Thrust

Sea Level, Static

Rolling Takeoff (M=0.25; SL)

Top-of-Climb (M=0.78; 35kft)

25,000

24,400

5,685

23,370

17,500

4,960

27,260

19,000

5,000

TSFC, lb/(lb·h)

Sea Level, Static

Rolling Takeoff (M=0.25; SL)

Top-of-Climb (M=0.78; 35kft)

0.344

0.480

0.610

0.257

0.352

0.494

0.158

0.229

0.428

Thrust-to-Weight

Sea Level, Static

Rolling Takeoff (M=0.25; SL)

Top-of-Climb (M=0.78; 35kft)

3.16

3.09

0.72

3.53

2.64

0.75

2.96

2.06

0.54

LTO Fuel Flow (kg/s)

Take-off

Climb Out

Approach

Idle

1.053

0.880

0.319

0.128

0.728

0.602

0.192

0.067

0.521

0.427

0.139

0.046

LTO NOX Emission Index (g/kg)

Take-off

Climb Out

Approach

Idle

26.5

22.3

8.9

4.7

16.5

10.7

9.0

5.0

14.0

9.0

9.0

5.0

LTO NOX, Dp/Foo (g/kN) 56.20 22.00 12.15
17

 

a) Geared turbofan. b) Geared open rotor. 

Figure 9. Advanced engine architectures (not to scale).

4.2 Aircraft Sizing and Performance 

The open rotor engine described above was combined with NASA’s ASAT-or airframe model to 

determine overall aircraft performance for the open rotor configuration. The advanced FPR=1.5 geared 

turbofan was also combined with the ASAT-re airframe model for a comparison configuration. Each 

configuration was sized to meet the same mission requirements: 162 passengers (32,400 lb), 3250 nm 

range, cruise Mach of 0.78. Wing area and engine thrust were optimized to meet the mission requirements 

with minimum gross weight, subject to constraints such as takeoff field length, second segment climb 

gradient, approach speed, landing field length, missed approach climb gradient, rate-of-climb at initial 

cruise altitude, and wing fuel volume. The characteristics of the resulting vehicles for each of the 

advanced engines are summarized in Table 2. Also included for comparison are results for the CSAT-re 

baseline technology airframe combined with the NASA V2525-D5-like engine model and sized for the 

same mission requirements. This vehicle is referred to as the “1990s Technology Baseline.” Weight, fuel 

consumption, and NOX emission results are also presented graphically in Figure 10.
18

Table 2. Aircraft Sizing Results 

1990s Tech. 

Baseline

ASAT-re

FPR=1.5 

ASAT-or

Geared OR

OEW, lb 94,450 79,650 87,800

Mission Fuel, lb 49,160 35,800 31,050

Payload Weight, lb 32,400 32,400 32,400

Ramp Weight, lb 176,000 147,850 151,270

Wing Area, ft21530 1240 1250

Wing Loading, lb/ft2115 119 121

Thrust (SLS), lb 25,190 23,075 26,900

Thrust-to-Weight (takeoff) 0.286 0.312 0.356

Takeoff field length, ft 7000 7000 6260

Landing field length, ft 5800 5940 6010

~Cruise Range Factor, nm

(velocity*(L/D)/TSFC)

12,670 14,730 17,170

Block Fuel, lb 41,550 30,390 26,710

Total NOX, lb 292 205 216

LTO NOX, lb per cycle 27.8 9.9 6.4

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

Operating Empty WeightGross Weight Block Fuel Total NOx LTO NOx Relative Values

1990's Technology Baseline Adv. Geared Turbofan Geared Open Rotor

Figure 10. Advanced vehicle weight, fuel, and emissions relative to 1990s technology baseline. 
19

As evident in Table 2 and Figure 10, both the advanced geared turbofan and advanced open rotor 

configurations have large weight, fuel consumption, and NOX emissions benefits over the 1990s 

technology baseline configuration. Empty weight reductions for the advanced configurations are 

relatively modest, mainly reflecting the benefits of a composite airframe and the cascading effects of a 

reduction in fuel weight. For the advanced open rotor vehicle, this benefit is offset to some extent by the 

much heavier engine. As a result, the open rotor configuration has higher empty weight and gross weight 

than the advanced geared turbofan configuration. Relative to the 1990s baseline, fuel consumption 

benefits are ~27% for the advanced geared turbofan and ~36% for the advanced open rotor. These 

reductions are from a combination of advanced airframe technology, advanced engine core technology, 

and advanced engine architectures. The reductions in total NOX emissions compared to the baseline are 

around 30% for both advanced configurations. Even though the open rotor configuration burns less fuel 

than the geared turbofan configuration, it emits more total NOX. The average cruise NOX EI is 5.1 for the 

advanced turbofan and 7.2 for the open rotor engine. Both engines have advanced low NOX combustors, 

but the combustion temperature is higher in the open rotor engine. LTO NOX emissions are greatly 

reduced for both advanced configurations. The LTO NOX emissions presented in Table 2 and Figure 10

are the estimated total NOX emissions produced during the landing-takeoff cycle. This value is derived by 

multiplying the ICAO Dp/Foo parameter by the total engine thrust. LTO NOX emitted by the open rotor 

advanced configuration is less than one-fourth of the amount emitted by the baseline. The advanced 

geared turbofan configuration has extremely low LTO NOX emissions as well, slightly more than onethird of the baseline value. 

To determine the true fuel consumption benefits of the open rotor architecture alone, comparison needs 

to be made to an equivalent technology turbofan configuration. The block fuel consumption of the 

advanced geared turbofan and advanced open rotor configurations are compared in Figure 11 for missions 

of various distances (all with 32,400 lb of payload). At the design point of 3250 nm, the open rotor 

architecture reduces fuel consumption 12% compared to an equivalent technology geared turbofan. This 

fuel consumption benefit is much less than the 20-25% reduction benefit often found in studies of the 

1980s. One reason for this lower benefit is the high cruise Mach number that is being pursued with 

modern open rotor engines. Modern blade design techniques have reduced the loss in efficiency 

associated with high Mach rotors, but the advantage of open rotors over ducted turbofans is still greater at 

the lower Mach numbers which were usually assumed in the earlier studies. A second reason for the 

smaller benefit in the current study is the ultra-high bypass ratio of the comparison geared turbofan. The 

efficiency benefits of the higher effective bypass ratio of the open rotor are less when the bypass ratio of 

the ducted turbofan engine is ~14 as in the current study instead of ~6 as in past studies. Finally, most of 

the past studies were focused on the benefits of open rotor propulsion on short range missions. As shown 

in Figure 11, the fuel consumption benefits of the open rotor engine increase at lower ranges. At a range 

of 500 nm, the reduction in fuel consumption for the open rotor grows to 18%. Table 3 compares the fuel 

consumption by mission segment for the open rotor and geared turbofan configurations on the 500 nm 

and 3250 nm missions. The fuel consumption benefits of the open rotor are the lowest during cruise, 

~10%, whereas fuel consumption reduction approaches 30% during other mission segments. For the short 

range mission, where cruise fuel is only about one-third of the total fuel use, the overall combined benefit 

is an 18% reduction in fuel use. (Note this particular mission profile is flown with optimum cruise 

altitude, leading to a climb mission segment that is almost as long as the cruise segment.) For the 3250 

nm mission, where cruise fuel accounts for over 85% of the total fuel use, the open rotor engine overall 

benefit drops to 12%. 
20

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

1.1

500 nm 1000 nm 1500 nm 3250 nm Relative Block Fuel

Mission Distance

Adv. GTF Geared OR

Figure 11. Block fuel comparison for advanced configurations. 

Table 3. Fuel Consumption (lb) by Mission Segment 

500 nm Mission 3250 nm Mission

Mission Segment Adv. Geared TF Adv. Open Rotor Adv. Geared TF Adv. Open Rotor

Taxi Out 142 102 (-28%) 142 102 (-28%)

Takeoff 395 284 (-28%) 395 284 (-28%)

Climb 2602 2045 (-21%) 2983 2353 (-21%)

Cruise 1786 1643 (-8%) 26,366 23,601 (-10%)

Descent 304 224 (-26%) 304 223 (-27%)

Approach 126 91 (-28%) 126 91 (-28%)

Taxi In 79 57 (-28%) 79 57 (-28%)

TOTAL 5,435 4,444 (-18%) 30,396 26,709 (-12%)

4.3 Certification Noise 

4.3.1 Trajectory Analysis 

Takeoff and approach trajectories and engine throttle settings must be modeled properly to correctly 

compute certification noise. An open rotor engine’s thrust characteristics differ from the performance of 

turbofans, and a trajectory calculation is necessary to capture these effects. This is in contrast to the 

simplified certification noise estimates often made during the conceptual phase of aircraft design, where 

fixed trajectories and throttle settings are assumed based on previously-collected data from other 

representative aircraft. Trajectory data computed for the open rotor airplane are shown in Figures 12 and

13. Aircraft parameters (altitude above field elevation (AFE), airspeed, Mach number, and angle of 

attack) are shown in Figure 12 and engine parameters (net thrust per engine, thrust fraction, advance ratio, 
21

and blade pitch angle) are shown in Figure 13. The trajectory is presented as a single operation with both 

takeoff and landing data shown simultaneously. For presentation purposes, the touchdown point on 

landing is coincident with the point of brake release on takeoff. The triangular markers on each plot 

denote the three noise certification measurement locations. The lateral observer is located on the lateral 

sideline at the distance from brake release where peak lateral noise is predicted. 

0.0

0.5

1.0

1.5

2.0

2.5

-15 -10 -5 0 5 10 15 20 25 30

Altitude, 1000 ft AFE

Distance from Brake Release, 1000 ft

0

50

100

150

200

250

-15 -10 -5 0 5 10 15 20 25 30

True Airspeed, ktas

Distance from Brake Release, 1000 ft

 (a) Altitude. (b) True airspeed. 

0.0

0.1

0.2

0.3

0.4

-15 -10 -5 0 5 10 15 20 25 30

Flight Mach

Distance from Brake Release, 1000 ft

0

2

4

6

8

10

12

14

16

-15 -10 -5 0 5 10 15 20 25 30

Angle of Attack, deg.

Distance from Brake Release, 1000 ft

 (c) Flight Mach number. (d) Aircraft angle of attack. 

 Figure 12. Takeoff and approach trajectory, aircraft parameters. 
22

0

5

10

15

20

25

30

-15 -10 -5 0 5 10 15 20 25 30

Net Thrust per Engine, 1000 lb

Distance from Brake Release, 1000 ft

0.0

0.2

0.4

0.6

0.8

1.0

1.2

-15 -10 -5 0 5 10 15 20 25 30

Thrust / Maximum Thrust

Distance from Brake Release, 1000 ft

(a) True net thrust per engine. (b) Thrust fraction. 

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-15 -10 -5 0 5 10 15 20 25 30

Front Rotor Advance Ratio

Distance from Brake Release, 1000 ft

25

30

35

40

45

-15 -10 -5 0 5 10 15 20 25 30

Front Rotor Pitch Angle, deg

Distance from Brake Release, 1000 ft

(c) Front rotor advance ratio. (d) Front rotor blade pitch angle. 

Figure 13. Takeoff and approach trajectory, engine parameters. 

4.3.2 Noise at Observers 

The EPNL database described in Section 3.3.4 was used to establish reference system noise levels at 

each observer. This process is illustrated in Figure 14 for the flyover observer. In the chart on the left, the 

isolated rotor flyover EPNLs at zero inflow angle are correlated with the front rotor tangential tip speed 

and the scaled thrust per engine. In the tunnel experiments, the tip speed was varied while blade pitch 

angles were held constant so that ranges of thrust could be investigated. But in practice, the front rotor 

would be operated at its constant design tip speed, even as the engine is throttled. (The blade pitch angles 

are flattened by a pitch-change mechanism as power is reduced.) In the chart on the right, the same 

EPNLs are plotted against the scaled thrust per engine. In this chart, a regression line is created using the 

locus of minimum EPNLs vs. thrust per engine. In both charts, the thrusts are scaled as described in the 

Appendix. Either chart may be used to determine the reference EPNL, but correlating the locus of 

minimum EPNLs against thrust is simpler since the influence of tip speed is removed. The reference 

system flyover EPNL (81.7 EPNdB) is at the intersection of the airplane cutback thrust per engine 

(11,270 lbs via the trajectory assessment) and the regression line. A similar process is used to determine 
23

the reference system EPNLs for the approach and lateral observers. The reference system noise levels are 

adjusted for rotor inflow angle effects and pylon effects as discussed in Section 3.3.5. Such “spreadsheet 

adjustments” are undesirable – a direct calculation of EPNL is preferred – but accurate rotor source noise 

prediction methods that include these effects are currently lacking. 

0 5000 10000 15000 20000 25000

Total Scaled Net Thrust per Engine, lb

1

2

3

4

5

6

7

8

9

1 0

81.7

Pitch Setting:

Cutback thrust level

75

80

85

90

95

100

400 500 600 700 800 900

Flyover EPNL, EPNdB

Tangential Tip Speed, ft/s

1

2

3

4

5

6

7

8

9

1 0

Pitch Setting:

10,000 lb

15,000 lb

20,000 lb

Thrust per Engine:

Figure 14. Reference system EPNL for the flyover observer.

The computed lateral, flyover, and approach EPNLs for the NASA open rotor vehicle are shown in 

Table 4 along with the corresponding Stage 3 limits for an airplane having a maximum takeoff gross 

weight of 151,300 lbs. For Stage 4 noise certification, the maximum noise level permitted at an individual 

measurement point is the same as the Stage 3 maximum level. However, Stage 4 aircraft are required to 

have a margin to Stage 3 of at least 2 EPNdB for any two measurement points combined, and they are 

required to have a cumulative margin to Stage 3 (all three measurement points combined) of not less than 

10 EPNdB. Additionally, “trading” EPNL margins between measurement points is no longer permitted. 

As seen from Table 4, the noise levels of an airplane using a modern rotor design would meet Stage 4 

requirements with a cumulative margin of 12.6 EPNdB. 

Table 4. Effective Perceived Noise Levels (EPNdB) for a 151,300 lb Aircraft 

Approach Lateral Flyover Cumulative

Reference Level 88.7 90.2 81.7 260.5

Inflow Effects 0.2 1.5 1.5 3.2

Pylon Effects 0.6 0.3 0.6 1.5

Overall 89.5 92.0 83.8 265.2

Stage 3 Rule 100.3 96.5 91.0 287.8

Stage 3 Margin -10.7 -4.5 -7.3 -22.6

Stage 4 Margin -12.6

5.0 Concluding Remarks 

An initial capability to assess and evaluate aircraft configurations with open rotor propulsion has been 

established in NASA. At present, this capability relies heavily on the availability of experimental data for 

existing rotor designs. Even though open rotor engines are heavy and require airframe accommodations 

such as additional noise insulation which further increase weight, when applied to an advanced singleaisle airframe the increase in propulsive efficiency results in large reductions in fuel consumption. The 

benefit relative to an equivalent technology geared turbofan configuration is 18% for short range 

missions, decreasing to 12% on long range missions. Because the open rotor engine has a higher cruise EI 
24

than the advanced geared turbofan, however, cruise NOX emissions are higher even though cruise fuel 

consumption is lower. Despite higher total NOX emissions, LTO NOX emissions for the open rotor 

configuration are 35% less than with an equivalent technology turbofan engine. Noise analysis for the 

modern low noise rotor design indicates a cumulative Stage 4 margin of 12.6 EPNdB. Development of 

low noise rotor designs continues beyond the rotors analyzed here, offering the potential for even better 

performance and lower noise. 
25

6.0 References 

1. Ethell, J. L.: Fuel Economy in Aviation. NASA SP-462, 1983. 

2. Bowles, M. D.: The “Apollo” of Aeronautics: NASA’s Aircraft Energy Efficiency Program 1973-1987. NASA 

SP-2009-574, 2010. 

3. Hager, R. D.; and Vrabel, D.: Advanced Turboprop Project. NASA SP-495, 1988. 

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

Space Technology, July 14, 2008. 

5. Norris, G.: Rig Rolling: Wing tunnel test imminent for Rolls’s open rotor rig. Aviation Week and Space 

Technology, July 14, 2008.

6. Conlon, J. A.; and Bowles, J. V.: Application of Advanced High Speed Turboprop Technology for Future Civil 

Short-Haul Transport Aircraft Design. AIAA 78-1487, 1978. 

7. Blythe, A. A.; and Smith, P.: Prospects and Problems of Advanced Open Rotors for Commercial Aircraft.

AIAA 85-1191, 1985. 

8. U.S. Code of Federal Regulations, Title 14, Chapter I, Part 36. Noise Standards: Aircraft Type and 

Airworthiness Certification. 

9. International Standards and Recommended Practices: Environmental Protection. Annex 16 to the Convention 

on International Civil Aviation, Vol. I: Aircraft Noise, 3rd ed., International Civil Aviation Organization 

(ICAO), Montreal, Canada, July 2008.

10. Current Market Outlook 2010-2029. Boeing Commercial Airplanes. 2010. 

11. Claus, R.W.; Evans, A.L.; Lytle, J.K., and Nichols, L.D.: Numerical Propulsion System Simulation, Computing 

Systems in Engineering, Vol. 2, No. 4, pp. 357-364, 1991. 

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

13. NPSS Reference Sheets Software Release: NPSS_1.6.5. 

14. Onat, E.; and Klees, G.: A Method to Estimate Weight and Dimensions of Large and Small Gas Turbine 

Engines. NASA CR 159481, 1979. 

15. 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, April 2004. 

16. 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. 

17. Guynn, M. D.; Berton, J. J.; Fisher, K. L.; Haller, W. J.; Tong, M. T.; and Thurman, D. R.: Engine Concept 

Study for an Advanced Single-Aisle Transport. NASA TM-2009-215784, August 2009. 
26

18. Hendricks, E. S.: Development of an Open Rotor Cycle Model in NPSS using a Multi-design Point Approach. 

GT2011-46694, ASME Turbo-Expo 2011, June 2011. 

19. Hendricks, E. S.; and Tong, M. T.: Performance and Weight Estimates for an Advanced Open Rotor Engine. 

AIAA-2012-3911, July 2012. 

20. Guynn, M. D.; Berton, J. J.; Fisher, K. L.; Haller, W. J.; Tong, M. T.; and Thurman, D. R.: Refined Exploration 

of Turbofan Design Options for an Advanced Single-Aisle Transport. NASA TM-2011-216883, January 2011.

21. Mann, S. A. E.; and Stuart, C. A.: Advanced Propulsion through the 1990s: An Airframer’s View. AIAA-85-

1192, 1985. 

22. McCullers, L.: Aircraft Configuration Optimization Including Optimized Flight Profiles. Proceedings of the 

Symposium on Recent Experiences in Multidisciplinary Analysis and Optimization, NASA CP 2327, April 

1984. 

23. MD-90-30 Airplane Characteristics for Airport Planning. MDC K9099, Boeing Commercial Airplanes, 

October 2002. 

24. Hopkins, H.: Powerfully Quiet: The McDonnell Douglas MD-90 is likely to meet any Stage 4 requirements, 

Flight International, October 26-November 1, 1994. 

25. Ardema, M. D.; Chambers, M. C.; Patron, A. P; Hahn, A. S.; Miura, H.; and Moore, M. D.: Analytical Fuselage 

and Wing Weight Estimation of Transport Aircraft. NASA TM 110392, May 1996. 

26. http://www.b737.org.uk/flightcontrols.htm. Accessed on 4/11/2011. 

27. Henne, P. A.: MD-90 Transport Aircraft Design. AIAA-89-2023, 1989. 

28. Whitfield, C. E.; Mani, R.; and Gliebe, P. R.: High Speed Turboprop Aeroacoustic Study, (Counterrotation), 

Volume I - Model Development. NASA CR 185241, July 1990. 

29. Whitfield, C. E.; Mani, R.; and Gliebe, P. R.: High Speed Turboprop Aeroacoustic Study, (Counterrotation), 

Volume II – Computer Programs. NASA CR 185242, July 1990.

30. Majjigi, M.; and Wojno, J.: Previous Open Rotor Noise Experience at GE. X-Noise Open Rotor Technology 

Seminar, Lausanne, Switzerland; March 18, 2011. 

31. NASA Press Release (Vol. 2, Issue 3, March 30, 2009): GE and NASA Partner on Open Rotor Engine Testing, 

http://www.nasa.gov/offices/oce/appel/ask-academy/issues/volume2/AA_2-3_F_ge_rotor.html. Accessed 

4/8/2011. 

32. Octave, Half-Octave, Third Octave Band Filters Intended for the Analysis of Sounds and Vibrations. 

International Electrotechnical Commission (IEC) Publication 225, 1966. 

33. Gillian, R.E.: Aircraft Noise Prediction Program User’s Manual. NASA TM-84486, 1983. 

34. Zorumski, W. E.: Aircraft Noise Prediction Program Theoretical Manual. NASA TM-83199, 1981, Parts 1 and 

2 (Currently maintained at NASA LaRC by the ANOPP team in electronic format and provided upon request; 

Latest revision: April, 2011). 
27

35. Rawls, J.: “ACD Module,” ANOPP Theoretical Manual, ver. 29, NASA Langley Research Center, Hampton, 

VA, 2011. 

36. Emmerling, J.J.; Kazin, S.B.; and Matta, R.K.: Core Engine Noise Control Program. Volume III, Supplement 1 

- Prediction Methods. FAA-RD-74-125, III-I, Mar. 1976. (Available from DTIC as AD A030 376) 

37. Stone, J.R.; Krejsa, E.A.; Clark, B.J.; and Berton, J.J.: Jet Noise Modeling for Suppressed and Unsuppressed 

Aircraft in Simulated Flight. NASA TM-2009-215524, March 2009.

38. Fink, M.R.: Airframe Noise Prediction Method. FAA-RD-77-29, March 1977. 

39. U.S. Code of Federal Regulations, Title 14, Chapter I, Part 25. Airworthiness Standards: Transport Category 

Airplanes. 

40. Standard Values of Atmospheric Absorption as a Function of Temperature and Humidity, SAE Aerospace 

Recommended Practice 866, 1964. 

41. Putnam, T.: Review of Aircraft Noise Propagation, NASA TM-X-56033, 1975. 

42. Delaney, M.; and Bazley, E.: Acoustical Properties of Fibrous Absorbent Materials, Applied Acoustics, Vol. 3, 

No 2, Apr 1970, pp 105-116.

43. Prediction Method for Lateral Attenuation of Airplane Noise During Takeoff and Landing, SAE Aerospace 

Information Report 1751, March 1981. 

44. Environmental Technical Manual on the use of Procedures in the Noise Certification of Aircraft, Document 

9501, 3rd ed., International Civil Aviation Organization (ICAO), Montreal, Canada, 2004. 

45. U.S. Code of Federal Regulations, Title 14, Chapter I, Part 34. Fuel Venting and Exhaust Emission 

Requirements for Turbine Engine Powered Airplanes. 
28

Appendix: Open Rotor Thrust Accounting and Scaling 

To make the best use of scale model open rotor acoustic data, a method for scaling measured thrust 

levels of open rotor test articles to appropriate sizes and flight conditions is needed. Thrust performance, 

power requirements, and acoustic spectra of scale model open rotor blade sets were measured in the 

NASA Glenn 9x15-foot Low-Speed Wind Tunnel. The open rotor propulsion rig’s data acquisition 

system provides model-scale rotor thrusts corrected to standard ambient conditions. In order for correct 

rotor noise levels to be matched to the airplane’s thrust demand at each noise certification condition, the 

thrust levels measured in the wind tunnel experiments must be scaled to full size and to the flight 

conditions determined by the aircraft trajectory calculations. This thrust-noise matching procedure 

requires that the thrust levels measured by the open rotor propulsion rig be scaled to the thrust that an 

actual full-size open rotor propulsion system would produce in flight. 

Dimensional Analysis 

A dimensional analysis is often used to gain insight to the behavior of physical systems. Consider a 

propeller blade of fixed geometry (i.e., the blade airfoil stack and pitch angle are fixed). It is reasonable to 

assume that the thrust (F) acting on the blade is a function of the following physical properties: the 

propeller diameter (D), its rotational speed (n), airspeed (uo), as well as the ambient fluid’s density (), 

bulk elastic modulus (K), dynamic viscosity (), static pressure (P), and the acceleration due to gravity 

(g). Other dimensional analyses of propellers have included additional (or fewer) properties, but these 

parameters are more than enough to determine a reasonable set of dependencies. In a dimensional 

analysis, the force acting on a rotor blade is assumed to take the form of a function: 

( )

21 3 a4 aaaa 8765

o

aa a

  gPKunDfF

Forcing dimensional homogeneity in the properties of mass (M), length (L), and time (T), the 

exponents ai may be related as: 

][ [( ])()()()()()()()

2 a1 1 a2 3 a3 1 a4 21 a5 11 a6 21 a7 2 a8 MLT TL ML LT ML T ML T ML T LT        



Equating terms in mass, length, and time, respectively, 

876542

8765431

7653

2 2222

31

1

aaaaaa

aaaaaaa

aaaa







Solving this system of equations for a1, a2 and a3 results in a meaningful functional relationship and 

the most sensible solution (in a Buckingham Pi dimensional analysis, D, n and  would be Buckingham’s 

repeating variables). After rearranging and substituting, a1 through a3 may be written as: 
29

3 765

2 87654

1 87654

1

2222

2224

aaaa

aaaaaa

aaaaaa







And so the rotor thrust is: 



























































































4 5 6 7 8

22 2 22 2

42

a a a a a

o

Dn

g

Dn

P

Dn nD

K

nD

u

fDnF









But the rotor wheel (tangential) tip speed, uT, is proportional to nD, so 































































































2 2 2

42

,,,,

T T T T T

o

u

gD

u

P

u Du

K

u

u

fDnF 

 







Further, the square of the speed of sound in a medium is K/, so the second collection of terms in the 

brackets is the tangential tip Mach number, MT. Also, a Reynolds number (Re) based on the tangential tip 

speed and with characteristic length D may be defined for the third collection of terms. In addition, the 

fourth collection of terms is expressed as the cavitation number (Ca), and the last terms are expressed as 

the Froude number (Fr). 

Finally, the dimensionless term uo/uT = uo/nD is defined as the propeller advance ratio, J, resulting in 

 MJfDnF FrRe Ca T

,,,,

42

 

Fluid cavitation and wave effects are generally applicable only to marine propeller performance, so for 

our purposes Fr and Ca are omitted. The propeller thrust coefficient is thus defined as 

  MJf Re

Dn

F

CT T, ,

42





A similar dimensional analysis may be performed for propeller power (resulting in the propeller power 

coefficient), but it is unrelated to this discussion of thrust scaling. 

CT remains relatively constant for geometrically similar propellers (i.e., ones in which the airfoil stack 

definitions, blade pitch angles and blade counts are identical, but may differ in scale), and for 

aerodynamically similar subsonic propellers (i.e., ones that advance, or corkscrew through the air at the 

same rate J). For minor excursions in altitude and airspeed within the takeoff regime (where 

compressibility and Reynolds effects are small, and provided J is constant), the flow patterns around the 

propeller are similar and F/n

2D4

remains a constant. 

Thrust Scaling in Similitude Conditions 

A propeller that maintains geometric and aerodynamic similarity throughout the takeoff and landing 

flight regime must (by definition) maintain a constant advance ratio. Thus, the tangential tip speed 

changes at the same rate as the airspeed. Substituting the definitions for the advance ratio (uo/nD) and the 
30

square of the sound speed (a

2

= P/ for a perfect gas, where is the specific heat ratio) into the thrust 

equation, 

2

2

2

2 2

2

42













 











 











 

J

DM

PCD

J

M

aCD

J

u

CDnCF

o

T

o

T

o

T T    

where Mo is the flight Mach number. Therefore, thrust scales by: 

22 F ~ PMo D

The simplest scaling is one where airspeed and tangential tip speed are fixed, and there are changes 

only in size and altitude. In this case, the thrust scales by: 

2

~ DF

Thrust Scaling in Non-Similitude Conditions 

In practice, variable-pitch (constant-speed) propellers operate in flight conditions that are decidedly 

not similar. For practical engine operational and performance reasons, the rotor tangential tip speed (and 

shaft speed) is held constant. Rotor blade pitch angle is varied as flight conditions change to provide the 

thrust required. As the airspeed increases from zero at brake release to climbout velocities of well over 

150 kts, the advance ratio increases as well. Blade pitch angles must increase with airspeed to maintain 

peak efficiency. With a constant-speed propeller, scaling thrusts from the wind tunnel Mach number 

(0.20) to the flight Mach number (often in excess of 0.26) using PM2D

2

 scaling is a very poor practice. 

For example, in going from the wind tunnel condition to the flyover observer flight condition, the thrust 

multiplier resulting from the Mach-square ratio is as high as 1.8. Without a commensurate increase in 

rotor tip speed, the rotor thrust would be drastically overpredicted. 

Ideally, if there were sufficient wind tunnel data collected at a variety of low-speed flight Mach 

numbers, a simple empirical thrust correction factor model could be developed. However, nearly all of the 

open rotor data were collected at a wind tunnel Mach number of 0.20. Therefore, instead, we seek a 

simple thrust correction multiplier that performs reasonably well for minor excursions in flight conditions 

in the takeoff regime. 

No scaling approach is perfect in non-similitude cases. Something reasonable in the absence of data is 

desired. From incompressible propeller actuator disk theory, it can be shown that the thrust is proportional 

to the difference in squares of the exit and entry flow velocities, ue

2

– uo

2

: 

2

22

oe

Disk

uu

AF



 

where ADisk is the disk area of the propeller. If changes in ue

2

 – uo

2

 are minor for the practical operation 

of a propeller throughout the takeoff regime, the rotor thrust is proportional to ambient density and 

diameter squared: 

2

~ DF

Density ratio scaling is attractive and intuitive since it reacts properly to changes in both altitude and 
31

ambient temperature. Scaling thrust with other altitude-dependent parameters is also possible. One 

alternative is the static pressure ratio. However, the static pressure ratio changes only with altitude (it is 

insensitive to changes in ambient temperature). Since thrusts must be scaled from wind tunnel conditions 

(corrected to international standard atmospheric standard conditions) to warmer standard acoustic day 

conditions (ISA+18°F), pressure ratio scaling is undesirable. 

However, ue

2

 – uo

2

 may not truly be constant, so a validation of non-similar density ratio scaling is in 

order. To examine the validity of density ratio scaling in the takeoff regime, a Monte Carlo simulation 

was conducted using propeller performance data for a commercially-available general aviation propeller. 

The reference point from which all other cases were scaled was standard sea level conditions at 130 kts. 

Altitude, airspeed, and ambient temperature were randomly varied from the reference condition to 

maximum values of 2500 ft in altitude, 150 kts in airspeed, and ISA+18°F in temperature. Propeller shaft 

speed was held constant while the shaft power was allowed to vary directly with pressure and inversely 

with the square root of temperature (constant corrected shaft power). Thrusts and blade pitch angles were 

outcomes. This analysis replicates the problem of scaling data from wind tunnel conditions to conditions 

found throughout the takeoff regime. 

The results of the Monte Carlo simulation are shown in the figure below. Thrusts obtained using 

density ratio scaling are indicated by the blue symbols, and thrusts obtained using pressure ratio scaling 

are indicated by the red symbols. Thrusts scaled from the reference condition using the density ratio 

compare reasonably well to the actual propeller thrust data. Density ratio scaling results in a smaller 

average error (about one percent) than pressure ratio scaling (about four percent). Positive error means the 

scaled thrust is greater than the actual thrust, so this analysis indicates that the scaling process will almost 

always overpredict the actual thrust. If the scaled thrusts are overpredicted, the process shown in Section 

4.3.2 may lead to underprediction of the aircraft noise levels. 

-15

-10

-5

0

5

10

15

370 380 390 400 410 420 430

Scaled Thrust Error, %

Actual Thrust, lb

Pressure ratio scaling

Density ratio scaling
32

Thrust Scaling 

The following thrust scaling relationship was used for this study: 

2

,

, 1 8

,, , 

















































  



Rig

ScaleFull

S L ISA

FISAFlight

AloneRotor

Rotor RamNozz

ScaleFullTotalN RigRotor D

D

F

F

F F





where 

FN,Tot,Full-Scale Full scale, total propulsion net thrust per engine 

FRotor,Rig Model scale, open rotor corrected thrust per engine (measured) 

FRotor+Nozz-Ram Total net thrust (rotor thrust plus core nozzle thrust minus core ram drag) 

FRotor Alone Rotor thrust alone 

Flight, ISA+18F Density at flight altitude, standard acoustic day (ISA+18°F) 

SL,ISA Density at sea level, standard day (0.002377 sl/ft3) 

(DFull-Scale/DRig)

2

 Area scale factor (square of the linear scale factor, model scale to full scale) 

The ratio FRotor+Nozz-Ram / FRotor Alone accounts for the small amount of thrust generated by the core nozzle 

and for the core ram drag. This ratio is determined from the analytical cycle model of the engine. In the 

model, the pressure exiting the power turbine at the aerodynamic design point at altitude is set so that the 

core nozzle pressure ratio at sea level is about 1.05. The “nozzle assist” generates a small amount of the 

total thrust near sea level. For NASA’s open rotor propulsion system near sea level, 

FRotor+Nozz-Ram / FRotor Alone = 1.027. 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Performance and Environmental Assessment of an Advanced Aircraft With 

Open Rotor Propulsion

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Guynn, Mark D.; Berton, Jeffrey J.; Haller, William J.; Hendricks, Eric; 

Tong, Michael T.

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

NASA Langley Research Center

Hampton, VA 23681-2199

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-20189

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 01

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

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

14. ABSTRACT

Application of high speed, advanced turboprops, or “propfans,” to transonic transport aircraft received significant attention during the 1970s and 1980s when 

fuel efficiency was the driving focus of aeronautical research. Unfortunately, after fuel prices declined sharply there was no longer sufficient motivation to 

continue maturing this technology. Recent volatility in fuel prices and increasing concern for aviation’s environmental impact, however, have renewed interest 

in unducted, open rotor propulsion. Because of the renewed interest in open rotor propulsion, the lack of publicly available up-to-date studies assessing its 

benefits, and NASA’s focus on reducing fuel consumption, a preliminary aircraft system level study on open rotor propulsion was initiated to inform decisions 

concerning research in this area. New analysis processes were established to assess the characteristics of open rotor aircraft. These processes were then used to 

assess the performance, noise, and emissions characteristics of an advanced, single-aisle aircraft using open rotor propulsion. The results of this initial study 

indicate open rotor engines have the potential to provide significant reductions in fuel consumption and landing-takeoff cycle NOX emissions. Noise analysis 

of the study configuration indicates that an open rotor aircraft in the single-aisle class would be able to meet current noise regulations with margin. 

15. SUBJECT TERMS

Aircraft design; Aircraft efficiency; Aircraft noise; Engine design; Open rotor propulsion

18. NUMBER

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42

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(443) 757-5802

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561581.02.08.07.42.03 

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2012-217772

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01- 10 - 2012