November 2011 

NASA/CR–2011-217303

Blended Wing Body Concept Development with 

Open Rotor Engine Integration

David M. Pitera, Mark DeHaan, Derrell Brown, Ronald T. Kawai, Steve Hollowell, Peter 

Camacho, David Bruns, and Blaine K. Rawden 

The Boeing Company, Huntington Beach, California 
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November 2011 

NASA/CR–2011-217303

Blended Wing Body Concept Development with 

Open Rotor Engine Intergration 

David M. Pitera, Mark DeHaan, Derrell Brown, Ronald T. Kawai, Steve Hollowell, Peter 

Camacho, David Bruns, and Blaine K. Rawden 

The Boeing Company, Huntington Beach, California 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 
1

1.1 Table of Contents

1.1 Table of Contents .................................................................................................... 1

1.2 Table of Figures ...................................................................................................... 2

1.3 Table of Tables ........................................................................................................ 3

2 Introduction ................................................................................................................... 4 

3 Requirements - Contract Task (3.1) .............................................................................. 4 

4 Configuration .............................................................................................................. 12

4.1 Wing-Body Development– Contract task (3.2) ...................................................... 12

4.2 Vertical Tail ............................................................................................................ 18

4.3 Nacelle Geometry .................................................................................................. 20

5 Propulsion Airframe Integration at Cruise, Unpowered ............................................. 22 

5.1 Overflow CFD Code .............................................................................................. 22

5.2 Isolated Nacelle ..................................................................................................... 23

5.3 Wing-Body ............................................................................................................. 24

5.4 Wing-Body-Nacelle ................................................................................................ 25

6 Propulsion Airframe Integration at Cruise, Powered .................................................. 33 

6.1 Method for Power .................................................................................................. 33

6.2 Wing-Body-Nacelle Configuration ......................................................................... 35

6.3 Adding Vertical Tails .............................................................................................. 35

6.4 Aligning the Nacelles ............................................................................................. 37

6.5 Wing-Body-Nacelle Configuration ......................................................................... 38

6.6 Adding Pylons ....................................................................................................... 40

6.7 High Speed Drag ................................................................................................... 42

7 Propulsion Airframe Integration, Low Speed, Powered ............................................. 43 

7.1 Low Speed Flight Conditions ................................................................................. 43

7.2 Wing Body ............................................................................................................. 43

7.3 De-rated Thrust ..................................................................................................... 49

7.4 Deflected Elevon ................................................................................................... 52

7.5 Thrust Drag Bookkeeping ...................................................................................... 55

8 Rotor Flow Field ......................................................................................................... 58

9 Concept Design Data Summary .................................................................................. 64 

9.1 Final Configuration ................................................................................................ 64

9.2 Drag Build-up ........................................................................................................ 65

9.3 Weight and Balance .............................................................................................. 69

9.4 Propulsion Characteristics ..................................................................................... 69

9.5 Performance .......................................................................................................... 70

10 Conclusions and Recommendations ........................................................................... 73 
2

11 References ................................................................................................................... 75 

1.2 Table of Figures 

Figure 1. BWB ERA-224 Upper Deck Cargo Arrangement. ................................................... 6

Figure 2. Payload Density versus Range. .............................................................................. 9

Figure 3. ERA-224 Point of Departure Configuration ........................................................... 12

Figure 4. Wing-Body Planform Changes .............................................................................. 13

Figure 5. Pressure Distributions for Different Wing Body Configurations ............................ 16

Figure 6. Spanload for different Wing Body Configurtaions ................................................. 16

Figure 7. 3-View Drawing Based on Final Optimized Geometry .......................................... 17

Figure 8. Aerodynamic Efficiency ......................................................................................... 17

Figure 9. Nacelle Geometry received from NASA GRC ....................................................... 21

Figure 10 . Definition of Inlet Cowl and Internal Lip .............................................................. 21

Figure 11. Baseline and Alternate Nacelle ........................................................................... 22

Figure 12. Isolated Baseline Nacelle Mach 0.80 .................................................................. 23

Figure 13. Isolated Alternate Nacelle at Mach 0.80 ............................................................. 24

Figure 14. Baseline Wing-Body Overflow Solution at M=0.80 ............................................ 25

Figure 15. Initial Wing-Body-Nacelle Solution ...................................................................... 26

Figure 16. CFD solution with a rotor tip to body spacing of 18 inches ................................. 27

Figure 17. Wing-Body Flow Field at the Centerline Nacelle Location .................................. 27

Figure 18. Effect of Nacelle Placement at Cruise ................................................................ 28

Figure 19. Low Drag Position of the Nacelles ...................................................................... 29

Figure 20. Preferred Stagger Arrangement for Low Noise ................................................... 30

Figure 21. Wing-Body-Nacelle, 3ft rotor tip to body spacing, unpowered ............................ 31

Figure 22. Wing-Body-Nacelle, 3ft Spacing, Mach number in horizontal off body plane .... 32

Figure 23. Wing-Body-Nacelle, 3 ft Spacing, Mach number in vertical off body plane ........ 32

Figure 24. Pressure Jump through an Actuator Disk on the Isolated Baseline Nacelle ....... 33

Figure 25. Velocity through an Actuator Disk on the Isolated Baseline Nacelle .................. 34

Figure 26. Mach Number through an Actuator Disk on the Isolated Baseline Nacelle ........ 34

Figure 27. Wing-Body-Nacelle with power and Mach number in off body plane .................. 35

Figure 28. Unpowered Overflow Solution with Baseline Nacelles Unaligned ...................... 36

Figure 29. Relative Position of the Engine Rotors and the Vertical Tails ............................. 37

Figure 30. Overflow Solution, Concept Design Data Config, Unpowered & Powered ......... 38

Figure 31. Drag Rise with Mach number from Overflow, Powered and Unpowered ............ 39

Figure 32. Final Configuration off body flow field in a horizontal plane ................................ 39

Figure 33. Overflow Solution for the Alternate Nacelle aligned with local flow .................... 40

Figure 34. Overflow Solutions with Pylons for the Baseline Nacelle .................................... 41

Figure 35. Overflow solutions with pylons for the baseline and alternate nacelles .............. 42

Figure 36. Power off and Power on Overflow Solutions at Stall ........................................... 44

Figure 37. Contracting Streamtubes Caused by High Power .............................................. 45

Figure 38. Severe Adverse Pressure Gradient Caused by High Power .............................. 45

Figure 39. Similar Separation Patterns from Different Methods for Simulating Power ........ 46

Figure 40. Separation Pattern with the Center Engine Forward .......................................... 47

Figure 41. Separation Pattern with a 3 ft Clearance between Rotor Tip and Body .............. 48

Figure 42. Reduced Separation at Higher Speed ................................................................ 49

Figure 43. Final Configuration, Power off with Closed Nacelles .......................................... 50

Figure 44. Final Configuration, 100% Thrust with Capped Nacelles .................................... 51
3

Figure 45. Final Configuration, 90% Thrust with Capped Nacelles ...................................... 51

Figure 46. Final Configuration, 80% Thrust with Capped Nacelles ...................................... 52

Figure 47. Powered Low Speed Configuration, Elevon at 0.0 Degrees ............................... 53

Figure 48. Powered Low Speed Configuration, Elevon at 10 Degrees TEU ........................ 53

Figure 49. Powered Low Speed Configuration, Elevon at 20 Degrees TEU ........................ 54

Figure 50. Powered Low Speed Configuration, Elevon Effectiveness ................................. 54

Figure 51, Isolated Nacelle Solution, Unpowered ................................................................ 56

Figure 52. Isolated Nacelle Solution, Powered .................................................................... 56

Figure 53. Powered and Unpowered Trimmed Drag Polars ................................................ 57

Figure 54. Elevator Required for Trim .................................................................................. 58

Figure 55. Circumferential Flow in the Forward Rotor Disk at Cruise .................................. 60

Figure 56. Mach Number in the Forward Rotor Disk at Cruise ............................................ 61

Figure 57. Flow Angle in the Forward Rotor Disk at Cruise ................................................. 61

Figure 58. Total Pressure in the Forward Rotor Disk at Cruise ........................................... 62

Figure 59. Circumferential Velocity in the Forward Rotor Disk at Takeoff Climb Out .......... 62

Figure 60. Mach Number in the Front Rotor Disk at Takeoff Climb Out .............................. 63

Figure 61. Flow Angle in the Front Rotor Disk at Takeoff Climb Out ................................... 63

Figure 62. Total Pressure in the Front Rotor Disk at Takeoff Climb Out .............................. 64

Figure 63. Final 3-View Drawing .......................................................................................... 65

Figure 64. Cruise Drag Bookkeeping ................................................................................... 66

Figure 65. Detailed High Speed Drag Build-Up ................................................................... 67

Figure 66. Hybrid Laminar Flow Control Assumptions ......................................................... 68

Figure 67. Riblet Assumptions ............................................................................................. 68

Figure 68. Thrust Trade for Minimum Fuel ........................................................................... 71

Figure 69. Results of Thrust Trade ...................................................................................... 72

1.3 Table of Tables 

Table 1. Container-Only Tare Weight and Internal Volume. .................................................. 7

Table 2. Pallet-Only Tare Weight and Internal Volume. ......................................................... 7

Table 3. Half-and-Half Mix of Containers and Pallets Tare Weight and Internal Volume. ..... 8

Table 4. Requirements. ........................................................................................................ 11

Table 5. List of Constraints for the MDOPT Optimization of the Wing Body Geometry ....... 14

Table 6. Vertical Tail Sizing .................................................................................................. 19

Table 7. Nacelle Rigging ...................................................................................................... 37

Table 8. Group Weight Statement ........................................................................................ 69

Table 9. OREIO Performance Sizing Constraints ................................................................ 70
4

2 Introduction

The BWB concept has been studied extensively due to the potential it has to 

significantly reduce fuel burn, emissions and noise levels compared to current and 

future conventional tube-with-wing subsonic transport configurations. The BWB 

concept has been studied with a variety of propulsion options, including podded 

nacelles, embedded engines, geared fans and open rotors. Open rotor technology 

has the potential to provide the greatest fuel burn reduction relative to all other 

potential propulsion options for the BWB. In addition, the BWB offers the potential 

for noise shielding, thus ameliorating one of the main drawbacks of the open rotor. 

However, open rotor integration with the BWB has not been studied as extensively 

as the podded or embedded engine options. The main purpose of this study is to 

perform a systems analysis of a BWB open rotor concept at the conceptual design 

level. This concept will be utilized by NASA to estimate overall noise and fuel burn 

performance, utilizing recent test data. The configuration is intended for entry into 

service (EIS) in 2025. The study was performed from the fall of 2010 to summer of 

2011 by the Boeing Research and Technology organization in Huntington Beach, 

CA. 

This study will also investigate the problem of Propulsion Airframe Integration (PAI) 

due to the installation of an open rotor configuration on a BWB. Open rotor engines 

have unique problems relative to turbo fans. The rotors are open, exposed to flow 

conditions outside of the engine. The flow field that the rotors are immersed in may 

be higher than the free stream flow and it may not be uniform, both of these 

characteristics could increase noise and decrease performance. The rotors 

sometimes cause changes in the flow conditions imposed on aircraft surfaces. At 

high power conditions such as takeoff and climb out, the stream tube of air that goes 

through the rotors contracts rapidly causing the boundary layer on the body upper 

surface to go through an adverse pressure gradient which could result with 

separated airflow. The BWB / Open Rotor configuration must be designed to 

mitigate these problems in order to reach fuel burn and noise goals.

 

The title of the project was long and cumbersome; an acronym was developed within 

Boeing to quickly refer to the project. The project was commonly referred to as 

OREIO, which stands for Open Rotor Engine Integration On a BWB. 

3 Requirements - Contract Task (3.1)

The purpose of the requirements is to support a design effort of a realistic airplane 

that supports further study by NASA. The requirements are based primarily on the 

requirements in development for the NASA/Boeing Environmentally Responsible 

Aviation (ERA) study of Blended Wing Body (BWB) and conventional airliners and 

freighters - specifically for the Preferred System Concept portion of the study. The 

hope is that common requirements will yield similar airplanes in both studies and 
5

that the in-depth analysis performed in this study will be valuable to the ERA study 

and its successors. It is recognized that the ERA requirements will evolve over time.

The following is intended to clarify how we arrived at the requirements. Only a few 

key requirements are spelled out for the ERA study. Other requirements are 

generally chosen to be representative of similar, existing airplanes to enhance the 

realism of the study.

Detailed requirements are shown in Table 4 at the end of this section. 

Payload

A cargo payload is chosen instead of a passenger payload. In most ways, a cargo 

payload is simpler than a passenger payload and permits the design of a realistic 

airplane with less concentration on the payload. This permits increased effort on 

propulsion-airframe integration – the object of this study.

We have chosen the point-of-departure (POD) airplane for the related ERA study 

effort as the starting point for this study as well. This airplane originated as an 

airliner, so despite the freighter-only payload, the airplane is configured to be a good 

airliner as well. It features a main passenger deck and a lower, central cargo hold.

The preliminary design of the POD airplane, presently called the ERA-224, includes 

the selection of cargo unit load devices (ULD). These are similar to existing ULD 

designs, but are not the same in that their dimensions (particularly height) are 

tailored for the BWB passenger cabin and lower hold dimensions.

The cargo arrangement is shown in Figure 1. There are 40 total positions of which 

22 are on the main deck. 18 primary ULDs 96” x 125” x 84” high are oriented 

widthwise in the three cargo bays. These are shown in yellow. Two additional ULDs 

88” x 125” with a tapered height from 82” to 62” fit into the forward end of the two 

outboard bays, as shown in cyan. 

In the lower hold, 18 rectangular ULDs 60.4” long x 61.5” wide by 46” high fit in two 

columns into the center bay. An alternative arrangement for the lower hold is 9 

widthwise pallets 60.4” long by 125” wide by 46” high.
6

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Tare weights for unit load devices are estimated from data presented in Reference 

1. Tares are estimated for two types of ULDs – all-containers and all-pallets with 

netting. The tare weights are substantial, especially for the containers.

Internal volumes of each of the container types is estimated based on the ratio of 

internal volume reported in Reference 1 compared to the volume of the ULD 

envelope (length x width x height). This ratio is approximately 92%.

Estimated tare weights of the ULDs in container-only form are shown in Table 1.

Total container tare weight is estimated to be 18,119 lb. Estimated tare weights of 

the ULDs in pallet-only form are shown in Table 2. Total tare weight is 60% lighter 

at 7228 lb. 

In practice, a mix of ULD types is used. Container loads are less expensive to build 

and more secure but more expensive to fly. Pallets are labor-intensive to build but 

are considerably lighter in tare weight. For the purpose of selecting a tare weight, a 

50-50 mix of containers and pallets is assumed. Results from this mix are shown in 

Table 3 with a total tare weight of 12,673 lb. Total cargo volume of the 50-50 mix, 

including an assumed 400 ft3 of bulk cargo is 13,689 ft3 as shown in Table 3. 
7

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Containers Only

Tare Weight

M1 Container reduced to 125" x 96" x 84" 15051.4

LD9 Container mod to 82/62" tapered height 1019.4

Lower Deck Containers 2048.4

Total tare weight (lb) 18,119

Container Internal Volume

M1 Container reduced to 125" x 96" x 84" 10868

LD9 Container mod to 82/62" tapered height 833

Lower Deck Containers 1638

Bulk Cargo (assumed) (ft^3) 400

Total internal volume (ft^3) 13,738

Assumed net cargo weight (lb) 100,000

Net cargo density (lb/ft^3) 7.279

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Pallets w/ Nets Only

Tare Weight

Main deck primary pallets w/ nets 5183.0

Main deck secondary pallets w/ nets 469.1

Lower deck pallets w/ nets 1575.6

Total tare weight (lb) 7,228

Container Internal Volume

Main deck primary pallets w/ nets 10733

Main deck secondary pallets w/ nets 843

Lower deck pallets w/ nets 1664

Bulk Cargo (assumed) (ft^3) 400

Total internal volume (ft^3) 13,641

Assumed net cargo weight (lb) 100,000

Net cargo density (lb/ft^3) 7.331
8

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50-50 Containers and Pallets

Tare Weight

Main deck primary 10117

Main deck secondary 744

Lower deck pallets 1812

Total tare weight (lb) 12,673

Container Internal Volume

Main deck primary 10800

Main deck secondary 838

Lower deck pallets 1651

Bulk Cargo (assumed) (ft^3) 400

Total internal volume (ft^3) 13,689

Assumed net cargo weight (lb) 100,000

Net cargo density (lb/ft^3) 7.305

NASA has specified a 100,000 lb cargo payload both for the ERA study and for the 

BWB Open Rotor Integration study. Initially it was not clear if this weight was to 

include or exclude the ULD tare weight. This choice influences the total payload of 

the airplane and the density of the cargo. NASA has now clarified the payload 

specification; the tare weight is included in the 100,000 lbs.

A more detailed study of NASA requirements shows that a BWB freighter of the 

ERA-224 configuration works well with at total (net + tare) payload of 100,000 lb. 

Although this payload yields a net cargo density of only 6.4 lb/ft3, the range at this 

payload is unusually long. When fuel is traded for payload, practical combinations of 

range and payload are achieved. A cargo density of 10.0 lb/ft3 yields an estimated 

range of about 3000 nm. This density is at the high end of commercial freighter 

densities while the range is at the low end, so this airplane nicely covers the typical 

range and density spectrum of commercial freighters. For comparison, the cargo 

density of the 767-300F is 7.2 lb/ft3 with a range of 3256 nm and the 777F is 9.9 

lb/ft3 at 4885 nm. The approximate relationship between the BWB, 767F and 777F 

payload densities and ranges is shown in Figure 2. 
9

Payload Density and Range

0

2

4

6

8

10

12

0 1000 2000 3000 4000 5000 6000 7000

Range (nm)

Payload Density (lb/ft^3)

BWB

777F

767F

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Performance

Performance requirements are again based on the NASA ERA requirements. The 

primary exception is that the cruise speed is reduced from Mach 0.85 to Mach 0.80. 

This reduction is to accommodate open rotor propulsion. Mach 0.80 is selected 

based on the consensus of the NASA-Boeing team.

Performance of the airplane will be estimated using Boeing’s standard advanced 

design assumptions as used in the recent NASA studies. These include:

x Performance is governed by Federal Aviation Regulations (FARs) for 

commercial transports.

x Boeing’s mission rules are used. These rules are significantly more efficient 

than current practice but are not as aggressive as those proposed for next 

generation airspace. These rules pertain to climb and descent profiles, hold 

times, taxi times, etc.

x A cruise-climb is assumed. This is representative of future airspace rules.

x A thrust reserve in cruise sufficient to provide at least 100 ft/min climb rate is 

required.

Airport Compatibility

Airport compatibility is addressed in three requirements. The load classification 

number (LCN) describes the load imparted by the landing gear on the airport 

surfaces and in turn defines the construction of these airport surfaces. A greater 
10

number indicates that a stronger surface is needed. The LCN is specified as being 

less than or equal to 90. 90 is a typical value for larger long range airliners. In 

advanced design, LCN is addressed by selecting the diameter and spacing of the 

main gear tires according to their loading.

Wingspan is constrained to 65 meters (213.25 ft). This is the limit for ICAO Code E 

airports, a reasonable limit for this capacity BWB that, unconstrained, optimizes near 

this span. Wing folding is not permitted in the interest of study simplicity although it 

may be entertained in future studies if the 65 meter limit proves to be a significant 

constraint. 

 

Detailed Requirements

Detailed requirements are listed in Table 4 below.

Measure of Merit

The initial configuration of the airplane will address noise with configuration features 

that may include canted inboard verticals for engine noise shielding, a forward 

engine location and extended and broadened aft fuselage for additional engine noise 

shielding and perhaps an outer wing free of leading edge devices to reduce slat 

noise. Once these features are selected and implemented, their refinement will be 

based primarily on drag considerations. Refinement of noise characteristics is 

beyond the scope of this study but may be addressed in the future.

Evolution and sizing of the airplane will consider fuel consumption as the measure of 

merit. This measure has proven effective in recent NASA sponsored studies of 

BWBs. 

Conclusion

The requirements described are intended to result in a practical, certifiable design. 

They are also intended to result in an airplane similar to that which results from the 

related ERA study. 
11

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BWB-224 Comments

Payload:

Pallet Positions 40

Tare Weight lb 12,673

1/2 Containers and 1/2 

Pallets with Nets

Total Cargo Volume ft3 13,689

Design Payload Density lb/ft^3 6.379 Computed w/o tare weight

Design Payload w/o Tare lb 87,327

Design Payload w/ Tare lb 100,000

Max Payload Density 10.000

Max Payload w/o Tare 136,892

Max Payload w/ Tare 149,565

Limit Load Factor g's 2.5

Sizing Unit Load Device (ULD) Type 96 x 125 x 84

1/2 Containers and 1/2 

Pallets with Nets

Sizing ULD Chamfering FALSE

Primary ULD Type and Number (20) 96 x 125 x 84

Primary ULD Volume ft3 10,800

Reduced Primary ULD (2) 88 x 125 x 82>62

1/2 Containers and 1/2 

Pallets with Nets

Reduced Primary ULD Volume ft3 838

Lower Deck Payload Type and Number

(18) 60.4L x 61.5W x 46H

1/2 Containers and 1/2 

Pallets with Nets

Lower Deck Payload Volume ft3 1651

Bulk Payload ft3 400 Assumed

Crew 2

Performance:

Range with design payload nm 6500

Range with maximum payload Fallout

Ferry range nm Fallout

Critical Field Length (SL Std Day @ MTOGW) ft 10,500

Standard for field length FAR

Max Landing Weight % TBD

Landing Weight (MLW) shall 

be the maximum zero fuel 

weight (MZFW) plus the 

reserve fuel required for a 

design range mission with 

Approach Speed KCAS 155

Landing Field Length (SL Std @ MLW) 5200 (dry)

Initial Cruise Altitude ft 35,000

Minimum Cruise Mach Number 0.80

Maximum Sinkrate @ Landing fps 8 Advanced Landing System

Airport Compatibility:

Runway Load Classification Number LCN <= 90

90 typ for LR airliners; 777-

300ER = 100

Wingspan Constraint meters 65 meters (213.25 feet)

Folding Wings FALSE

Noise

Emphasize a low noise 

design - but no specific 

requirement in terms of 

dB.

Include noise-shielding 

configuration features such 

as canted inboard verticals, 

forward-mounted engines 

and a widened and lengthed 

aft centerbody.

Measure of Merit

Gross payload ton-nm per 

lb fuel burned

At design range. Aircraft 

acoustic features 

uncompromised.
12

4 Configuration

4.1 Wing-Body Development– Contract task (3.2)

The wing and body lines were developed from the ERA-224 configuration shown in 

Figure 3. The ERA-224 configuration is a close match for the requirements from the 

previous section with the exception of the cruise Mach number. The 0.85 cruise 

Mach number for the ERA was recognized as being too high for open rotors. The 

cruise Mach number specified for this study is 0.80. The wing body lines of the 

ERA-224 configuration had to be changed primarily by reducing the wing sweep to 

match the cruise Mach number. 

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The wing sweep was modified in 3 steps. First, a new sweep angle was determined 

based on sweep-thickness requirements for a conventional wing. Second, the body 

centerline chord was increased to allow more room for the propulsion system and to 

increase the body area aft of the engines for better noise shielding. Third, the 

aerodynamic characteristics were improved by using the MDOPT code to optimize 

the configuration. The three configurations are labeled ERA-080, ERA-080REV2, 

and ERA-080R2DLR3 and are shown in Figure 4. 

The centerline chord was increased mainly because it was felt that the upper surface 

of the body may have to be carved out or “area-ruled” to reduce the interference 

between the engines and the body. The volume removed for area-ruling was 

estimated to begin forward of the rear spar and there was not enough room at the 

rear spar to reduce the height of the body. The increased centerline chord length

relieved this problem. Also, by increasing the chord while maintaining the same 

9

CHARACTERISTICS

Wing Span (ft) 212.6

Wing Area (sf) 8,313

Length (ft) 115.5

WAG Fan Dia. Shown (in) 188

Thrust (lb) ~31,300

Cruise Mach Number 0.85

Total Swet (sf) 19,306

b2/Swet 2.341

Payload 20-96x125x84H

2-88x125x84H Shaped

18-LD-2-46

2-Bulk Cargo

223 PAX, 3C, Long Range Rules 

787-8 b=197.2 Swet=19,862

Wetted AR = 1.958

PAYLOAD:

Containers @ 92%E 7.74 lbs/cf – 95,930 lbs

Bulk Cargo @ 100%E 7.74 lbs/cf – 4,480 lbs

Total – 100,410 lbs Required – 100,000 lbs

FUEL:

402,685 lbs. available. 3X+ what is needed.

WAG’s: (as scaled from EET-612)

OEW (lbs) 223,000

Payload (lbs) 100,000

Fuel (lbs) 124,000

MTOGW (lbs) 447,000

T/W 0.21

Thrust (lbs) 93,870

Problematic to offset rotors more 

than shown – must resolve
13

thickness reduced the thickness to chord ratio which reduced the aerodynamic 

penalty for adding the extra wetted area.

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However, the increased centerline chord ultimately proved to be a problem and was 

removed for the final configuration. CFD analysis during the propulsion airframe 

integration task indicated that in order to position the nacelle and pylon in slower 

velocity air flow field they should be located as close to the trailing edge as possible. 

The increase in the centerline chord moved the trailing edge aft; the nacelle and 

pylon moved aft along with the trailing edge. The increased planform area and the 

aft movement of the engines increased the increased overall weight and resulted 

with a center of gravity too far aft for proper balance of the configuration. 

The optimization was done using MDOPT an aero and multidisciplinary constrained 

optimization that is based on a design of experiments. The CFD code that was 

coupled to the optimizer for this work was TLNS (Thin Layer Navier-Stokes). Only 

the wing and body were modeled for ease of gridding and speed of computation. A 

table of constraints used in MDOPT is contained in Table 5. The optimization also 

constrained the center of pressure between 39% and 40% of the MAC. Thrust 

effects were accounted for by assuming that thrust equals drag and that the moment 

arm for the thrust vector was 180” above the center of gravity. The optimization was 

done at a lift coefficient of 0.25 and an altitude of 35,000 ft.

After the MDOPT optimization was complete the pressure distribution was smoothed 

using CDISC, a CFD based inverse design tool coupled with a flow solver called 

CFL3D. The final geometry was then smoothed manually. 

ERA-224 to ERA-080

• Reduced sweep for reduced 

Mach

• M = 0.85 - 0.80

ERA-080 to ERA-080REV2

• Increased centerline chord

• Improved propulsion installation 

options and noise shielding

• Maintain Tmax not (T/C)max

GROSS PLANFORM AREAS

ERA-224 7967 FT2 

ERA-080 7967 FT2 

ERA-080REV2 8499 FT2 

ERA-080R2DLR3 8377 FT2 

ERA-080REV2 to ERA080R2DLR3

x MDOPT optimization with 

trim constraint

x CDISC design to smooth 

pressures for a smooth loft

x Manual geometry smoothing
14

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The optimized wing-body geometry has more sweep and more outboard loading 

than the original wing-body. These changes not only improved the lift-to-drag ratio 

for the configuration but also helped to balance the airplane at 39% of the MAC with

thrust effects from the engine mounted 180” above the upper surface of the body.

The increased sweep is shown in Figure 4. The increased outboard loading is 

evident in the pressure distribution shown in Figure 5 and the spanload distribution 

shown in Figure 6. A 3-view was completed base on the final geometry after 

optimization and smoothing. The 3-view drawing is shown in Figure 7. 

The maximum lift-to-drag ratio for the final wing-body configuration is in line with 

other wing-body designs from CFL3D as shown in Figure 8. Note that the efficiency 

of the wing-body alone configurations will shift over to the left slightly when the extra 

wetted area of the nacelles, pylons and vertical tails are added.

x Optimization done on clean wing 

(no N/P or verticals)

x Maximize L/D subject to the 

center of pressure constraint 

39%MAC < C.P. < 40%MAC at 

CL=0.25/35KFT

x C.P. includes effect of thrust

x Thrust = Drag

x Moment arm assumed at 180” 

above wing surface at rotor 

location

x Thrust axis alignment considered 

secondary effect and ignored

Design Variable Table

 Lower Upper Station Final

 Index Type Bounds Bounds Location Value

1 Sweep -5 5 0 2.52

2 Taper Ratio -0.05 0.05 0 -0.013

3 Span -0.05 0.05 0 0.002

4 Twist -0.01 0.01 0.157 0.01

5 Twist -2 2 1 1.21

6 Max T/C Location -0.01 0.01 0.157 -0.006

7 Max T/C Location -0.1 0.1 0.6 -0.070

8 Max T/C Location -0.1 0.1 1 0.047

9 TE Deflection -3 3 0 -0.93

10 TE Deflection -3 3 0.157 -0.69

11 TE Deflection -0.1 0.1 0.235 0.06

12 Camber Splines -0.0001 0.0001 0.157 -0.00009

13 Camber Splines -0.0001 0.0001 0.157 -0.00001

14 Camber Splines -0.0001 0.0001 0.157 0.00003

15 Camber Splines -0.01 0.01 0.6 0.00422

16 Camber Splines -0.01 0.01 0.6 0.00773

17 Camber Splines -0.01 0.01 0.6 0.00930

18 Camber Splines -0.01 0.01 1 -0.00676

19 Camber Splines -0.01 0.01 1 0.00398

20 Camber Splines -0.01 0.01 1 0.00086

NOTES: TE hinge line at X/C=0.65

Camber spline knots at X/C=0.2, 0.5, 0.8
15

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)LJXUH6SDQORDGIRUGLIIHUHQW:LQJ%RG\&RQILJXUDWLRQV
16

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Aerodynamic Cruise Efficiency

14

16

18

20

22

24

26

28

30

1 1.2 1.4 1.6 1.8 2 2.2

Span / (Swet*Cf/0.003)^0.5

100% Implies Fully-Turbulent Max Potential, CDp = CDi, e = 1, Zero Trim, Zero Compressibility Drag

100%

80%

90%

70%

Recent BWB's

Current BWBs

Wing Alone CFL3D

(L/D)max

M=0.84-0.85

M=0.70-0.80

%

Early BWB's and 

Tube & Wing Aircraft

ERA-080

)LJXUH$HURG\QDPLF(IILFLHQF\

X

Y

2552 in.

879 in.

383 36 in

1460 in.

675.82 in.

17.36°

BWB – Open Rotor – R3

224 passengers

3 – Bays

2 – Decks
17

4.2 Vertical Tail

A revised vertical tail was developed for the CFD analysis. The new tail was 

necessary because the sweep on the vertical tail used for the original OREIO 

configuration is too low for an aircraft with a cruise Mach number of 0.80. The new 

tail is based on the area and taper ratio of the tail from the original OREIO 

configuration. The quarter chord sweep of the new tail was set at 35 deg which is 

appropriate for a cruise Mach number of 0.80 based on conventional sweep and 

thickness charts. The Mach capability of the tail must be greater than the wing to 

maintain control for all conditions. Standard symmetrical airfoils were used; they 

were twisted to get zero lift at cruise based on the flow field of the wing-body.

The revised vertical tail was also checked relative to rules of thumb from NASA 

based on results from the acoustic testing in the LSAF tunnel. Based on results 

from acoustic testing in the LSAF tunnel, NASA suggested sizing rules for the root 

chord and tip chord relative to the forward rotor diameter. The root chord should be 

about 1.25 times the front rotor diameter and the tip chord should be about 0.60 

times the front rotor diameter. The root chord and tip chord for the revised vertical 

tail are 1.13 and 0.48 times the front rotor diameter. The revised vertical tail is close 

to the recommended size for acoustics but slightly undersized.

The size of the vertical tail was also checked relative to requirements for stability and 

control. The OREIO configuration and the X-48C are very similar, and the X-48C

was evaluated in great detail. The X-48C tail volume should be a very good 

preliminary sizing requirement for OREIO. The tail volumes are shown in Table 6.

Two OREIO configurations are shown: the original OREIO configuration (Figure 7), 

and the final OREIO configuration (Figure 63) which will be discussed in Section 

10.1. The OREIO configurations have approximately the same tail volume as the 

X48-C, which implies that the OREIO tails are the correct size for stability and 

control. Therefore the vertical tail size was not increased. 

7DEOH9HUWLFDO7DLO6L]LQJ

Vertical Tail Sizing for Stability and Control

Match the Tail Volume of the X-48C

X-48C

Original 

OREIO

Configuration

Final OREIO

Configuration

Tail Arm 68.652 572.21 492.20

Tail Area 1.7 169.6 169.6

Wing MAC 90.797 887.9 824.3

Wing Area 100.48 8379.8 8002.5

Tail Volume 0.01279 0.01304 0.01265
18

The final geometry resulted with the height of the top of the front rotor approximately 

equal to the height of the top of the vertical tail. This is shown in Figure 29. The 

front rotor-to-body spacing shown in Figure 28 is 3 ft based on requirements for 

minimum air flow separation on the upper surface of the body at takeoff conditions 

explained in Section 8.2. 

4.3 Nacelle Geometry

The external nacelle geometry is based on the nacelle geometry received from 

NASA shown in Figure 9. The design is based on the GE UDF open rotor work from 

the 1980’s. The external cowl and inlet were not sufficiently defined so standard

design methods were used. The external nacelle cowl was defined based on a DAC 

3 shape. A two to one ellipse was used for the inlet lip. The geometry of the inlet 

cowl is shown in Figure 10.

The technology level of the rotors used by NASA is also based on 1980’s work. The 

latest technology open rotors generally have a higher blade count, increased 

clipping of the rear rotor diameter relative to the front rotor, and a lower thrust 

loading. The lower thrust loading (thrust per square foot of rotor area) for the latest 

technology rotors results in a larger diameter. The original 3-view drawing for the 

OREIO project was set up assuming the latest technology rotors with an 18” rotor tip 

to body spacing. The NASA rotors have a smaller diameter so when the NASA 

rotors were placed in the same engine location as the 3-view the rotor tip to body 

spacing was increased to 30”. 

An alternate nacelle was developed and used for some of the CFD solutions. The 

alternate nacelle was developed because CFD solution with the baseline nacelle

indicated high levels of interference between nacelles, with the body, and with the 

pylons when they were included. The alternate nacelle is shown in Figure 11. The 

baseline nacelle has a large diameter because it was designed for a Mach number 

of 0.72. The alternate nacelle has a smaller maximum diameter and therefore lower 

external velocities and less interference with other parts of the configuration. The 

diameter of the alternate nacelle is sufficient to enclose the engine, and provide the 

proper external flow characteristics around the inlet and cowl including low spillage 

drag. However, the alternate nacelle does not have the same shape or diameter 

near the rotor hubs and will therefore have different airflow characteristics in the 

rotor planes. The alternate nacelle has a constant diameter in front of the rotor 

planes. The baseline nacelle has a continuously varying diameter reaching a 

maximum at the end of the cowl and then reducing forward of the rotor planes. The 

alternate nacelle would have a different flow field at the rotors because it lacks the 

conditioning provided by the shape of the baseline nacelle. The baseline nacelle 

was not changed because of concern about changes to the rotor flow field.

However, some CFD solutions were produced with the alternate nacelle for 

comparison to the baseline configuration. 
19

)LJXUH1DFHOOH*HRPHWU\UHFHLYHGIURP1$6$*5&

)LJXUH'HILQLWLRQRI,QOHW&RZODQG,QWHUQDO/LS

HWB Open Rotor (8x8)Layout X=138.2

” 

X=0

45.1

77.7 28.7

X=45.0

X=85.0” X=118.8

X=155.3

12 5

X=106

X=112

46.0 36.0 40 153. 168.

X=82

X=204

21

19.4

26.8

X= -

X=143.6

BWB Open Rotor Integration Inlet Shape

R(throat) = 14.35

R(hl) = 16.979

X(start of throat) = -39.542

X(hl) = -44.8 X(end of cowl) = 45

Engine centerline

DAC 3 Shape

X

Y
20

)LJXUH%DVHOLQHDQG$OWHUQDWH1DFHOOH

5 Propulsion Airframe Integration at Cruise, Unpowered

5.1 Overflow CFD Code

The Overflow RANS (Reynolds Averaged Navier-Stokes) code was used for most of 

the CFD analysis. The Overflow code was selected because it has proven accuracy 

and Boeing in Huntington Beach has a large amount of experience with it. The 

accuracy of Overflow is well documented in a series of papers from the AIAA’s Drag 

Prediction Workshop. The Overflow code was also selected because it has a rotor 

model which has been used for helicopter and open rotor work on other projects. 

Overflow is a structured code that uses overset grids. The structural grid can take 

more time to prepare than an unstructured code, but Boeing has so much 

experience with it that many short cuts have been developed. Also the geometry for 

this project is rather simple. Many of the configuration changes involved moving the 

nacelles for configurations without a pylon. The grid preparation for these cases 

was a quick and easy process even when using a structured code. 

HWB Open Rotor (8x8)Layout 

X=138.

2”

X=0

45.1

” 

77.7 28.7

X=45.0

X=85.0” X=118.8

X=155.3

12.5

X=106.

X=112.

46.0 36.0 40.2 153.

5”

168.

8”

X=82.

X=204.

21.3

19.4

26.8

X= - 

X=143.6

Baseline Nacelle

Alternate Nacelle
21

5.2 Isolated Nacelle

Isolated Overflow solutions for the baseline and alternate nacelles were produced at 

0.80 and 0.84 Mach number. A mass flow plug was developed to set the mass flow 

ratio to 0.60 which is appropriate for this type of engine. The mass flow plug was 

taken from another application and scaled to get the correct mass flow ratio. The 

plug is shaped differently than what is shown in Figure 9 but the specific shape of 

the plug aft of the exit plane will have little impact on the solutions. The solution for 

Mach number 0.80 for the baseline nacelle is shown in Figure 12 and for the 

alternate nacelle in Figure 13. The pressure coefficient is shown in the lower right 

corner of each figure. The internal pressure coefficients are positive and are nearly 

identical for both the baseline and alternate nacelles. The external pressure 

coefficients are more negative for the baseline configuration indicating that the 

external velocities are higher. The biggest difference in the pressure coefficients 

occurs in the area where the alternate nacelle has a constant section. The alternate 

nacelle has less negative pressure coefficients in this area which indicates lower 

velocities. The lower velocities outside the alternate nacelle indicate that it should 

produce less interference drag than the baseline nacelle.

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE

BASELINE NACELLE AT M=0.80

)LJXUH,VRODWHG%DVHOLQH1DFHOOH0DFK
22

Boeing Research and Technology / Aero-Sciences

BOEING i t d k f B i M t C

ISOLATED NACELLE

ALTERNATE NACELLE AT M=0.80

)LJXUH,VRODWHG$OWHUQDWH1DFHOOHDW0DFK

5.3 Wing-Body

An Overflow solution for the baseline wing-body is shown in Figure 14. The wing 

body solution is very similar to the results produce by CFL3D during the wing-body 

development. 
23

Boeing Research and Technology / Aero-Sciences

WING ONLY 

M=0.80; CL=0.25

)LJXUH%DVHOLQH:LQJ%RG\2YHUIORZ6ROXWLRQDW0 

5.4 Wing-Body-Nacelle 

The engine integration work was started by evaluating the interference drag of the 

wing-body-nacelle configuration. Pylons were not included because of the 

complexity of gridding the geometry and the difficulty in developing an acceptable 

configuration. The interference caused by a pylon would be so large that it would be 

difficult to determine the effects of nacelle positioning. Vertical tails were not 

included for the same reason, although it was assumed that the interference due to 

the vertical tails would be less than the interference from the pylon. It was also 

recognized that vertical tails are critical for the noise shielding aspects of the 

configuration and would have to be evaluated before long. 

The wing, body, nacelle work started by evaluating the nacelles in the position 

shown on the 3-view drawing. The 3-view drawing shown in Figure 7 placed the 

nacelles side-by-side next and quite far forward of the trailing edge. The distance 

between the nacelles and the trailing edge was increased due to the trailing edge 

extension that was discussed in Section 5.1. The interference drag was high for this 

configuration. The Overflow solution for this configuration is shown in Figure 15; the 

interference drag is 8.1 counts. Note that the interference drag is calculated by 

subtracting the drag for the wing-body and three isolated nacelle solutions from the 

wing-body-nacelle solution. The total drag of the configuration is about 110 counts, 

so 8.1 counts is approximately 7% of the total drag. 
24

An indication of interference drag is evident in the pressure distribution. The 

pressure distribution for the isolated nacelle is shown in Figure 12. The pressure 

distribution for the installed nacelles is shown in Figure 15. The same scaling was 

used for the pressure coefficients in both figures. The pressure distributions for the 

installed case have regions of high velocity that are not found in the isolated 

pressure distributions. Higher velocities are indicated by the increased amount of

red on the nacelle surface for the installed case. Note that there are two legends 

shown in Figure 15. The legend on the lower left shows the pressure coefficients on 

the surface of the nacelles and aft body. The legend on the upper right shows the 

local Mach number in off body planes above and below each nacelle. A peak local 

Mach number near 1.2 is shown in the off-body planes indicating a moderate 

recompression shock between the nacelles and the body. The recompression shock 

is also evident on the body and on the sides of the nacelles. 

A second rotor tip to body spacing was also evaluated. The rotor tip to body spacing 

for the configuration in Figure 15 is 30 inches. This spacing resulted because the 

rotor provided by NASA has a smaller diameter than the rotor that was used for the 

original 3-view. This was discussed in Section 5.3. The desired rotor tip to body 

spacing is 18 inches. This spacing will lower the engines and reduce the nose down 

pitching moment due to the thrust axis. The Overflow solution with a spacing of 18 

inches is shown in Figure 16. The interference drag increased from 8.1 to 10.4 

counts when the rotor tip to body spacing was reduced from 30” to 18”. The 

increased shock strength on the body is evident under the nacelles.

Boeing Research and Technology / Aero-Sciences

NACELLE PLACEMENT STUDY

POWER OFF

NOMINAL POSITION #1 (Nacelles side by side, no stagger)

30 in propeller-to-wing clearance

)LJXUH,QLWLDO:LQJ%RG\1DFHOOH6ROXWLRQ
25

Boeing Research and Technology / Aero-Sciences

WING/NACELLE

INSTALLED BASELINE NACELLE

AT POSITION #2

)LJXUH&)'VROXWLRQZLWKDURWRUWLSWRERG\VSDFLQJRILQFKHV

Boeing Research and Technology / Aero-Sciences

WING ONLY 

Mach Contours at Centerline of Center Engine

)LJXUH:LQJ%RG\)ORZ)LHOGDWWKH&HQWHUOLQH1DFHOOH/RFDWLRQ

The interference drag was attacked in two ways. First, the flow field in which the 

nacelle was immersed was evaluated from the wing-body grid solution. This flow 
26

field, shown in Figure 17, indicated that the nacelle should be moved aft. There are 

strong velocity gradients over the body caused by its shape and the lift that it 

generates. From the front to the back of the imaginary nacelle the velocity 

decreases from Mach 0.85 to Mach 0.80. The nacelle was not actually in the 

solution; it is just shown for reference. The second method of reducing the 

interference drag was to stagger the nacelles. The original positioning of the 

nacelles side-by-side created a local maxima of cross sectional area that caused a 

spike in velocity. This maximum cross sectional area could be reduced by 

staggering the nacelles.

The interference drag was reduced by about two thirds by moving the nacelles aft 

and staggering them. The interference drag for several wing-body-nacelle 

combinations is shown in Figure 18. The first group of nacelle position changes was 

to move the center nacelle aft in 20” increments; the interference drag dropped 

quickly as the stagger increased. The stagger was stopped at 60” assuming that 

positioning the engines and blades at a significantly different fuselage station would 

not be good for acoustics or for integrating the engines to the back of the wing-body. 

)LJXUH(IIHFWRI1DFHOOH3ODFHPHQWDW&UXLVH

A second group of nacelle position changes consisted of just two configurations 

where all three nacelles were moved aft closer to the trailing edge. The interference 
27

drag was lower for the second group of configurations. One configuration had a 40” 

stagger and 2.7 counts of interference drag. The other had a 60” stagger and 2.4 

counts of interference (Figure 18). The Overflow solution for the 60” stagger is 

shown in Figure 19. 

Boeing Research and Technology / Aero-Sciences

NACELLE PLACEMET STUDY

POWER OFF

LOW DRAG POSITION (Center Engine Aft)

2 ft propeller-to-wing clearance

)LJXUH/RZ'UDJ3RVLWLRQRIWKH1DFHOOHV

The spacing between the rotor tip and the body was less for the second group of 

configurations. The nacelles in the first group were moved aft but not down. The 

nacelles in the second group were adjusted down when they were moved aft. The 

decreased spacing penalized the interference drag of the second group, but the 

interference drag of this group was still lower than the first group because all three 

nacelles were moved aft to slower air near the trailing edge.

The movement of the nacelles aft closer to the trailing edge of the body was limited 

by a criterion supplied by NASA; to ensure adequate noise shielding the middle of 

the forward and aft rotors should be at least one forward rotor diameter from the 

trailing edge.

The movement of the nacelles close to the trailing edge ruined the reasoning for 

adding the 80” trailing edge extension on the body. The extension moved the 

trailing edge of the body further aft than the original configuration but then the 

nacelles were moved aft to the lower velocities near the trailing edge. The additional 

structure associated with the 80” longer body and aft location of the nacelles created 

a weight and balance problem that was solved by reducing the extension by 80” for 

the Concept Design Data Summary (Section 10.1). 
28

The stagger of the nacelles was changed based on inputs from NASA for low noise. 

The center engine was staggered 60” forward of the outboard engines instead of 60” 

aft. This way the center engine would be further from the trailing edge and have 

more noise shielding and the outboard nacelles would be shielded by the vertical 

tails. Results from the Overflow solution for this inverted stagger configuration are 

shown in Figure 20. The interference drag increased slightly from 2.4 counts to 2.5 

counts.

Boeing Research and Technology / Aero-Sciences

NACELLE PLACEMET STUDY

POWER OFF

INVERTED STAGGER (Center Engine Forward)

2 ft propeller-to-wing clearance

)LJXUH3UHIHUUHG6WDJJHU$UUDQJHPHQWIRU/RZ1RLVH

The low speed evaluation in Section 8 also had an influence on the nacelle and 

pylon configuration. The Overflow solution at stall speed with high power resulted 

with a significant amount of separation on the upper surface of the body with a 24” 

separation between the rotor tip and the body. In order to reduce the amount of 

separation on the upper surface of the body, the nacelle was moved up to a 36” 

spacing. A high speed Overflow solution for this final wing-body-nacelle unpowered 

configuration is shown in Figure 21. A top view of the same solution with the 

addition of local Mach numbers interpolated on a horizontal plane that bisects the 

center engine is shown in Figure 22. There is a separate legend in Figure 22 for the 

pressure coefficients on the nacelles and the local Mach number on the horizontal 

plane. Mach numbers greater than 1.0 are evident between the nacelles. The high 

Mach number forward of the nacelles is due to the lifting wing-body. A side view 

with a similar vertical plane is shown in Figure 23. Once again the local Mach 

number plane bisects the center nacelle. High Mach numbers and a well defined 

shock are evident between the nacelle and the surface of the body. 
29

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ HIGH SPEED

M=0.80; CL=0.25; POWER-OFF

NACELLES: INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

CL = 0.25

CD = 100.5 cts.

• Drag is 0.4 counts lower 

than the configuration 

with a 2 ft clearance

)LJXUH:LQJ%RG\1DFHOOHIWURWRUWLSWRERG\VSDFLQJXQSRZHUHG
30

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ HIGH SPEED

M=0.80; CL=0.25; POWER-O FF

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

CD = 100.5 cts.

)LJXUH:LQJ%RG\1DFHOOHIW6SDFLQJ0DFKQXPEHULQKRUL]RQWDORIIERG\SODQH

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ HIGH SPEED

M=0.80; CL=0.25; POWER-O FF

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

CD = 100.5 cts.

Figure 23. Wing-Body-Nacelle, 3 ft Spacing, Mach number in vertical off body plane
31

6 Propulsion Airframe Integration at Cruise, Powered

6.1 Method for Power

Power was simulated in the OREIO study by two methods: an actuator disk, and a 

rotor model. The actuator disk is a standard method of applying a pressure jump 

through the disk where the thrust is equal to the pressure jump multiplied by the disk 

area. In the case of this dual rotor geometry, both discs were assumed to produce 

the same amount of power per square foot. Since the forward rotor has more area, it 

produces more thrust. An example of an isolated nacelle with the actuator disk used 

for power is shown in the following Figures. The pressure jump across the disk is 

shown in Figure 24. The increase in velocity though the disk in shown in Figure 25.

The velocity increase is smooth even though the pressure is discontinuous. The 

Mach number though the disk is shown in Figure 26. The Mach number jumps 

because the pressure jump causes a temperature jump which affects Mach number.

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE@ HIGH SPEED

M=0.80; ACTUATOR DISK

TEMPERATURE THROUGH FWD ROTOR

)LJXUH3UHVVXUH-XPSWKURXJKDQ$FWXDWRU'LVNRQWKH,VRODWHG%DVHOLQH1DFHOOH

The rotor model is a relatively simple method that captures general rotor effects 

without the need to model moving blades. The rotor model adds momentum 

sources in the direction normal to the rotor plane to simulate thrust and tangentially 

to simulate blade drag. The result of the rotor model is temporally averaged thrust 

and swirl imparted to the flow field as momentum. The rotor model was attempted in 

some cases; however the rotor model needs more development for high speed (high 
32

Mach number) cases. The actuator disk was the primary method for simulating 

power for the OREIO project. 

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE@ HIGH SPEED

M=0.80; ACTUATOR DISK

VELOCITY MAGNITUDE THROUGH FWD ROTOR

)LJXUH9HORFLW\WKURXJKDQ$FWXDWRU'LVNRQWKH,VRODWHG%DVHOLQH1DFHOOH

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE@ HIGH SPEED

M=0.80; ACTUATOR DISK

MACH NUMBER THROUGH FWD ROTOR

)LJXUH0DFK1XPEHUWKURXJKDQ$FWXDWRU'LVNRQWKH,VRODWHG%DVHOLQH1DFHOOH
33

6.2 Wing-Body-Nacelle Configuration

The powered Overflow solution for the wing-body-nacelle configuration has a minor 

increase in drag due to power. The solution is shown in Figure 27. The drag is 0.8 

counts higher than the unpowered solution shown in Figure 22. By comparing the 

amount of red shown in Figure 22 and Figure 27, it is apparent that the shock 

strength between the nacelles has increased.

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ HIGH SPEED

M=0.80; CL=0.25; ACTUATO R DISK

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

CD = 101.8 cts.

)LJXUH:LQJ%RG\1DFHOOHZLWKSRZHUDQG0DFKQXPEHULQRIIERG\SODQH

6.3 Adding Vertical Tails

The revised vertical tail described in Section 5.2 was added to the configuration and 

twisted so that it would not lift at cruise. The new tail was placed in the same 

location as the old tail by matching the location of the quarter chord of the MAC for 

both tails. A twist distribution was developed by evaluating the flow-field of the 

powered wing-body-nacelle configuration. The twist angles were then checked by 

placing a floating vertical tail above the wing body, and finally by attaching the 

vertical tail and checking the lift at cruise. The intent was to have zero lift on the 

vertical tail at cruise. The floating vertical tails were used as a quick way to check 

for alignment without developing a complex grid that would be required by the 

intersection of the vertical tail and the body. A minor adjustment was made to the 

mid span airfoil twist after installation. The final rigging angles of the vertical tail are 

10 degrees nose out at the root and 4 degrees nose out at the tip.
34

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ HIGH SPEED

M=0.80; CL=0.25; POWER-OFF

NACELLES: INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

CL = 0.25

CD = 108.1 cts.

• Drag increase due to 

verticals is 7.6 counts

• Estimated drag of the 

verticals, using wetted 

area, skin friction, and 

form drag, is 3.3 counts

Forw ard vertical

)LJXUH8QSRZHUHG2YHUIORZ6ROXWLRQZLWK%DVHOLQH1DFHOOHV8QDOLJQHG

The initial Overflow solutions with the vertical tail had large amounts of interference 

drag. An unpowered solution is shown in Figure 28. The configuration drag was 

later reduced by aligning the nacelles as explained in Section 41.

The vertical tail was moved aft 40” to improve noise shielding for the outboard 

engines. NASA suggested to move the nacelles aft but did not recommend how far. 

A distance of 40” was used. The relative positions of the vertical tail and the engine 

rotors are shown in Figure 29. In the planform view before the 40” movement, the

rotors were very close to the trailing edge of the vertical tail. The planform view is 

improved with the vertical tails 40” farther aft. However, the side view is opposite; 

more of the rotors are hidden with the vertical tails in the forward position than in the 

aft position. These pictures would look different if real swept blade geometry was 

used. Another thing to consider is the direction of noise propagation from the rotors; 

the aft position may be better.
35

Boeing Research and Technology / Aero-Sciences

NACELLE ALIGNMENT

ORIGINAL ALIGNMENT / REVIS ED ALIGNMENT

ORIGINAL

NACELLE

ALIGNMENT

ALTERNATE 

NACELLES

CENTER: PITCH /TOE 3.49/0.00

OUTBD: PITCH/TOE 2.06/3.59

REVISED

NACELLE

ALIGNMENT

VERTICALS TRANSLATED 40 IN. AFT

BASELINE 

NACELLES

CENTER: PITCH /TOE 3.85/0.00

OUTBD: PITCH/TOE 2.48/4.00

)LJXUH5HODWLYH3RVLWLRQRIWKH(QJLQH5RWRUVDQGWKH9HUWLFDO7DLOV

6.4 Aligning the Nacelles

The nacelles were aligned normal to the cruise onset flow to reduce the amount of 

flow distortion into the rotors. It was assumed alignment was necessary to provide 

high propulsive efficiency, long rotor life, and low cabin noise. All three goals would 

be impacted if the rotors were misaligned and experiencing a loading cycle with 

each revolution. Pointing the thrust vector away from the direction of flight will also 

reduce cruise efficiency but it was assumed that subjecting the rotors to cyclic 

loading would be worse. These assumptions need to be confirmed by an engine 

manufacturer.

The pitch and toe angles required for alignment were larger than expected. The 

rigging angles for the final configuration are listed in Table 7.

7DEOH1DFHOOH5LJJLQJ

Nacelle Rigging 

Toe 

Angle Incidence Angle

Center Engine 0.00 3.85

Outboard Engine 2.48 4.00
36

Drag of the wing-body-nacelle-vertical tail configuration was reduced by aligning the 

nacelles. Aligning the nacelle is a powerful method to reduce drag. A significant

drag reduction occurred for aligning the nacelles for both the baseline and alternate 

nacelles, for powered and unpowered cases, and for cases using the rotor model or 

the actuator disk to simulate power. 

6.5 Wing-Body-Nacelle Configuration

The final configuration uses the baseline nacelles aligned as described in Section 

7.4 with the aft location for the vertical tails as described in Section 7.3. Overflow 

solutions for this configuration are shown in Figure 30. The interference drag due to 

power is only 0.5 counts. The drag rise predicted from Overflow is shown in Figure 

31. The drag rise is similar for both the powered and unpowered solutions. Drag 

divergence occurs at about a Mach number of 0.80. The largest difference in drag 

between the powered and unpowered cases is about 1.5 counts at low Mach 

numbers.

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ HIGH SPEED

M=0.80; CL=0.25; POWER OFF VS ON (ACTUATOR DISK)

NACELLES: ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

POWER OFF POWER ON

CD=106.2 cts. CD=106.7 cts.

)LJXUH2YHUIORZ6ROXWLRQ&RQFHSW'HVLJQ'DWD&RQILJ8QSRZHUHG	3RZHUHG
37

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ HIGH SPEED

DRAG RISE

DRAG RISE

CL=0.25

100.0

102.0

104.0

106.0

108.0

110.0

112.0

114.0

116.0

118.0

120.0

0.60 0.65 0.70 0.75 0.80 0.85

M

CD

POWER ON

POWER OFF

THRUST CONSTANT

)LJXUH'UDJ5LVHZLWK0DFKQXPEHUIURP2YHUIORZ3RZHUHGDQG8QSRZHUHG

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ HIGH SPEED

M=0.80; CL=0.25; ACTUATOR DISK

NACELLES: ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

)LJXUH)LQDO&RQILJXUDWLRQRIIERG\IORZILHOGLQDKRUL]RQWDOSODQH
38

The flow field between the nacelles and the vertical tail has regions of high Mach 

number but the magnitude is less than before the nacelles were aligned. The off 

body flow field in a horizontal plane is shown in Figure 32. There is less air flow over 

Mach 1.0 in Figure 32 than in Figure 27.

The alternate nacelle has less drag than the baseline nacelle. An Overflow solution 

for the alternate nacelle is shown in Figure 33. The drag is 3 counts lower with the 

alternate nacelle than with the baseline nacelle. In this case 3 counts of drag is very 

close to 3% of cruise drag making the alternate nacelle a clear winner. The problem 

is that the large diameter of the baseline nacelle is not designed for the high Mach 

numbers that exist above the wing-body. The baseline nacelle is designed to 

condition the flow for the rotors when flying isolated at Mach 0.72. It is not clear 

which nacelle has the best flow field for the rotors when installed. 

Boeing Research and Technology / Aero-Sciences

ALTERNATE ALIGNED NACELLES

PYLON-OFF

CD* = 103.7 cts.

* No T/D bookkeeping performed

D= 4.500 deg.

)LJXUH2YHUIORZ6ROXWLRQIRUWKH$OWHUQDWH1DFHOOHDOLJQHGZLWKORFDOIORZ

6.6 Adding Pylons

Pylons have been avoided so far in this study. The previous section reviewed the 

final configuration for the study; it did not have pylons. Several attempts were made 

to add pylons to the configuration near the end of the study. There was no effort to 

develop a final pylon that could be considered acceptable for this configuration. The 

quick looks were just to see how bad the problem was.

Two simple pylons were created. The pylons were created from a simple symmetric 

airfoil with a 12% thickness to chord ratio. One pylon has a 15 degree sweep and 
39

the other has a 30 deg sweep. The 15 degree swept pylon was used because it 

would be closer to what was drawn on the 3-view where the pylons are nearly 

vertical so that they can provide better support for the engine. The 30 degree swept 

pylon was used because it is closer to the sweep that would be required for a 12% 

thick wing at Mach 0.80. The outboard pylons were toed out 4 degrees to align them 

with the local onset flow.

The depth of the spars assumed for the original OREIO configuration in Figure 7 is 

20”. The pylons that were used for the CFD evaluation will only support spars that 

are about 12” deep.

The Overflow solutions with pylons have significant amounts of interference drag. 

The solutions are shown in Figure 34. The configuration with the 15 degree swept 

pylon has a total drag of 126.2 counts. The configuration with the 30 degree swept 

pylon has a total drag of 117.1 counts. This is compared to 106.7 counts without the 

pylons. There is a significant amount of separation for both configurations but it is 

certainly worse for the 15 degree swept pylon. The 30 degree pylon has a classic 

aerodynamic advantage with higher sweep which is more appropriate for the cruise 

Mach number, but it may also have had a smoother progression of cross sectional 

area through the aft body region. 

Boeing Research and Technology / Aero-Sciences

BOEINGi t d k f B i M t C

EFFECT OF PYLON SWEEP

BASELINE ALIGNED NACELLES

* No T/D bookkeeping performed

* *

)LJXUH2YHUIORZ6ROXWLRQVZLWK3\ORQVIRUWKH%DVHOLQH1DFHOOH

The alternate nacelle had less drag when coupled to the 30 degree swept pylon. 

The Overflow solution for the baseline and alternate nacelles with the 30 degree 
40

swept pylon is compared in Figure 35. The solution with the alternate nacelle has 

4.9 counts less drag.

Boeing Research and Technology / Aero-Sciences

EFFECT OF NACELLES

30 DEG. SWEPT PYLONS

* No T/D bookkeeping performed

* *

)LJXUH2YHUIORZVROXWLRQVZLWKS\ORQVIRUWKHEDVHOLQHDQGDOWHUQDWHQDFHOOHV

6.7 High Speed Drag

An estimate of the high speed interference drag was developed. The total drag from 

Overflow for the wing-body-nacelle configuration was 106.7 counts. The 

interference drag of the wing-body-nacelle configuration is about 5 counts, based on 

an estimate of the vertical tail drag and the isolated wing-body and isolated nacelle 

drag from Overflow. The interference for the installations with the pylon are much 

higher. A decision was made to include an interference drag level of 4.5 counts for 

the nacelle and pylon installation. This includes 3 counts of interference for the 

nacelles and 1.5 counts of interference for the pylons. The 4.5 count increment was

included in the total drag build up (Figure 65).
41

7 Propulsion Airframe Integration, Low Speed, Powered

7.1 Low Speed Flight Conditions 

One of the goals established early in the OREIO study was that there should be no 

separation near stall. It was assumed that at the beginning of a takeoff run, the high 

power though the rotors combined with the low speed would lead to upper surface 

separation; this would be acceptable as long as there was no separation when the 

aircraft was flying. Separated airflow near the rotors could cause a non-linear 

response from the center elevon. It would also increase drag; the engine would 

have to produce more thrust to compensate for the drag which would increase noise. 

To ensure that there was no separation present when flying, the slowest flying speed 

possible was evaluated, this is the stall speed. Since the stall speed has not been 

established for the OREIO configuration an assumption was made that the stall 

speed would be at about Mach 0.15. A leading edge device such as a slat or 

Krueger has not been defined either so a solution for the cruise wing at an angle-ofattack of 10 degrees was used. A second condition was also evaluated. The 

second condition represented the takeoff climb out speed. This was a faster 

condition that was just used for comparison with the stall speed case because it was 

difficult to eliminate all the separation from the stall speed case. The speed used for 

the takeoff climb out condition was Mach 0.20.

The thrust used for the Overflow evaluations came from the data provide by NASA. 

The thrust levels from NASA for Sea Level static and Mach 0.25 were 37,000 lbs 

and 29,000 lbs. The thrust levels for CFD were interpolated from this data. The 

thrust level for Mach 0.15 and 0.20 were 32,200 lbs and 30,600 lbs. 

7.2 Wing Body 

Powered Overflow solutions for the wing body had significant upper surface 

separation. Power off and power on Overflow solutions at stall are shown in Figure 

36. For this comparison the angle-of-attack for the power on solution was reduced 

so that both solutions were at the same lift coefficient. The amount of separation is 

considered unacceptable. The effectiveness of the elevon behind the engines will 

be reduced and may be non linear. 
42

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LOW SPEED

M=0.15; CL=0.48

)LJXUH3RZHURIIDQG3RZHURQ2YHUIORZ6ROXWLRQVDW6WDOO

The Overflow solution was evaluated to see if other characteristics of the flow field 

could help to understand the reason for the separation on the upper surface of the 

body. Streamlines were traced up and down stream of the rotor tips to develop 

stream tubes of airflow that go though the rotors. The stream tubes are shown in 

Figure 37. The area of the stream tubes decreases rapidly as the airflow is 

accelerated through the rotors. On the surface of the body the high power 

accelerates the airflow forward of the rotors, but then when the cross sectional area 

of the stream tubes decrease at the rotors, there is a large adverse pressure 

gradient. The pressure gradient on the upper surface of the body is shown in 

Figure 38. The adverse pressure gradient is causing the separation on the body.
43

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LOW SPEED

POWER-ON; M=0.15; CL=0.48

)LJXUH&RQWUDFWLQJ6WUHDPWXEHV&DXVHGE\+LJK3RZHU

)LJXUH6HYHUH$GYHUVH3UHVVXUH*UDGLHQW&DXVHGE\+LJK3RZHU
44

The upper surface separation was similar for two different methods of simulating 

power. The separation pattern from an Overflow solution using the rotor model was 

compared to the pattern from a solution using an actuator disk and is shown in 

Figure 39. There are some differences but the area that is separated is generally 

the same in both solutions. 

53

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LOW SPEED

M=0.15; D=10 DEG.

NACELLES: LOW DRAG POSITION

ROTOR TIP-TO-WING UPPER SURFACE 2 FT.

ACTUATOR DISK MODEL

OLD DISK GRIDS

ZERO MFR THRU CORE

ROTOR MODEL

)LJXUH6LPLODU6HSDUDWLRQ3DWWHUQVIURP'LIIHUHQW0HWKRGVIRU6LPXODWLQJ3RZHU

The Overflow solutions that were discussed so far were completed early in the study 

before the engine stagger was changed by moving the center engine forward. The 

separation pattern with the center engine staggered forward is different than what 

was shown so far; this pattern is shown in Figure 40. There is still severe separation 

but it might be slightly better than what is shown in Figure 39. 

In order to relieve the separation problem the engine was moved up so that there 

was a distance of 3 ft between the rotor tip and the body upper surface. The upper 

surface separation pattern with a 3 ft clearance is shown in Figure 41. The 

separation pattern has improved but was not completely eliminated. 

The decision to increase the distance between the rotor tip and the body was made 

after consulting with others on the project to determine if the stall condition was too 

severe and maybe an unrealistic case to evaluate. The conclusion was that it was 

not unrealistic. If this airplane is built and stall tested, it will be flown to stall at high 

power and all the control surfaces must have good effectiveness and a linear 
45

response. In order to have the proper response there cannot be a large amount of 

separation on the upper surface of the body. Flow control on the upper surface of 

the body is another option to reduce separation that was discussed. Flow control 

may be effective but is beyond the scope of this study. 

A second faster condition was evaluated to see how quickly the separation would 

decrease with speed. The condition selected is the takeoff climb out speed. It was 

selected because it is within the typical operation envelope of the aircraft. The stall 

speed is not within the operational envelope of the aircraft, only the flight test aircraft 

will be evaluated at stall. The takeoff climb out condition is at a speed of about 

Mach 0.20. The angle-of-attack would also be slightly lower, but the angle-of-attack 

of 10 degrees was maintained for consistency; the only change would be speed and 

thrust. The flow solution reflected the same engine and power setting. The thrust 

was only changed because of the typical lapse rate that would be experienced. The

upper body airflow patterns from the Overflow solutions from Mach 0.15 and Mach 

0.20 are shown in Figure 42. There is certainly less separation at the faster speed. 

It is difficult to say that there is no separation, there is no reversed flow, but there are 

streamlines that converge in a manner that may indicate that off body vortices are 

present. Further configuration changes due to low speed separation were stopped 

after seeing these results. Some additional work was done to help understand the 

problem, but there were no configuration changes. The configuration with a distance 

of 3 ft between the rotor tip and the body upper surface became the baseline for the 

remainder of the study. The configuration being evaluated at cruise conditions was

also changed to the 3 ft clearance.

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LOW SPEED

M=0.15; D=10 DEG.

NACELLES: INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 2 FT.

• 2 ft Clearance

• Inverted stagger results 

with a similar amount of 

separation but with a 

different pattern

)LJXUH6HSDUDWLRQ3DWWHUQZLWKWKH&HQWHU(QJLQH)RUZDUG
46

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LOW SPEED

M=0.15; D=10 DEG.

NACELLES: INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

• 3 ft clearance

• Inverted stagger has a 

similar amount of 

separation as the opposite 

stagger but with a different 

pattern

)LJXUH6HSDUDWLRQ3DWWHUQZLWKDIW&OHDUDQFHEHWZHHQ5RWRU7LSDQG%RG\
47

Boeing Research and Technology / Aero-Sciences

WING/NACELLE @ LO W SPEED

EFFEC T O F MACH; ALPHA=10 DEG.; POWER-O N

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

M=0.15 M=0.20

)LJXUH5HGXFHG6HSDUDWLRQDW+LJKHU6SHHG

7.3 De-rated Thrust

A brief study was done to see how quickly the upper surface separation problem 

improved with reduced thrust. The nacelles used for this study had capped off inlets 

and exits as explained in Section 8.5. Four Overflow solutions were completed at an 

angle-of-attack that was approximately 8 degrees but varied slightly so that a 

constant lift coefficient was maintained. The stall condition was used for this 

evaluation. The stall condition is at a Mach number of 0.15 and has a significant 

amount of separation on the upper surface of the body. The results are shown in 

Figure 43 through Figure 46. The first solution is unpowered and shows that the 

upper surface of the body has attached airflow at this condition. The first solution 

with power show significant separation. As power is reduced to 90% and then 80% 

of full takeoff power the separation improves but is not eliminated. Something else 

will have to be changed on the configuration, if the separation must be eliminated to 

provide the proper control effectiveness from the center elevon. Flow control 

through suction or blowing is one possibility.
48

The aircraft performance resulting from de-rating the engine thrust was checked. 

The configuration can meet the takeoff field length requirement of 10,500 ft with a 

SLST of 33,870 lbs or a de-rate of about 90%.

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ LOW SPEED

M=0.15; CL=0.476; POWER-OFF

NACELLES: CLOSED/ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

)LJXUH)LQDO&RQILJXUDWLRQ3RZHURIIZLWK&ORVHG1DFHOOHV
49

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ LOW SPEED

M=0.15; CL=0.476; ACTUATOR DISK

NACELLES: CLOSED/ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

)LJXUH)LQDO&RQILJXUDWLRQ7KUXVWZLWK&DSSHG1DFHOOHV

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ LOW SPEED

M=0.15; CL=0.476; ACTUATOR DISK

NACELLES: CLOSED/ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

)LJXUH)LQDO&RQILJXUDWLRQ7KUXVWZLWK&DSSHG1DFHOOHV
50

Boeing Research and Technology / Aero-Sciences

WING/NACELLE/VERTICAL @ LOW SPEED

M=0.15; CL=0.476; ACTUATOR DISK

NACELLES: CLOSED/ALIGNED/INVERTED STAGGER

ROTOR TIP-TO-WING UPPER SURFACE 3 FT.

)LJXUH)LQDO&RQILJXUDWLRQ7KUXVWZLWK&DSSHG1DFHOOHV

7.4 Deflected Elevon

The control effectiveness of the center elevon was evaluated using the Overflow 

CFD code. The center elevon was deflected both trailing edge up and trailing edge 

down, at a Mach number of 0.15 and 0.20. The upper surface separation and the 

pitching moment change for each solution were evaluated. This work was 

completed before the nacelles were aligned. The alternate nacelle was used 

because at that time in the study it was felt that the alternate nacelle would become 

the primary nacelle. The nacelle alignment and shape will not have much of an 

effect on the results; the key flow feature is the large pressure jump through the 

rotors required to produce the thrust required for takeoff. Results are shown in 

Figure 47 through Figure 50. Flow separation is present at a Mach number of 0.20 

for the center elevon at deflection angles of 0.0, 10.0, and 20.0 degrees trailing edge 

up. At 0.0 degrees of deflection the amount of separation is small, similar to Figure 

42, but the amount of separation increases as the trailing edge is moved up. The

pitching moment provided by deflecting the center elevon is plotted in Figure 50.

The pitching moment is surprisingly linear for all three elevon deflection angles. The 

elevon also produces similar levels of pitching moment and at both Mach numbers 

and it has already been shown that there is more separation at the lower Mach 

number. The question of the effectiveness of the center elevon with power requires 

more investigation.
51

Boeing Research and Technology / Aero-Sciences

WING/ALT. NACELLE/VERTICAL @ LO W SPEED

M=0.2; ALPHA=8 DEG.; PO WER-O N

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

)LJXUH3RZHUHG/RZ6SHHG&RQILJXUDWLRQ(OHYRQDW'HJUHHV

Boeing Research and Technology / Aero-Sciences

WING/ALT. NACELLE/VERTICAL @ LO W SPEED

M=0.2; ALPHA=8 DEG.; ELEVO N -10; POWER-O N

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

)LJXUH3RZHUHG/RZ6SHHG&RQILJXUDWLRQ(OHYRQDW'HJUHHV7(8
52

Boeing Research and Technology / Aero-Sciences

WING/ALT. NACELLE/VERTICAL @ LO W SPEED

M=0.2; ALPHA=8 DEG.; ELEVO N -20; POWER-O N

NACELLES: INVERTED S TAGG ER

RO TO R TIP-TO -WING UPPER SURFACE 3 FT.

)LJXUH3RZHUHG/RZ6SHHG&RQILJXUDWLRQ(OHYRQDW'HJUHHV7(8

Boeing Research and Technology / Aero-Sciences

CONTROL SURFACE EFFECTIVENESS

• LOOK AT ELEVON DEFLECTIONS OF 0, 10, -20 AT M=0.15 AND M=0.20

• DETERMINE ELEVON EFFECTIVENESS ('CM/'Ge)

CM VS Ge

-0.1000

-0.0800

-0.0600

-0.0400

-0.0200

0.0000

0.0200

0.0400

0 -5 -10 -15 -20 -25

Ge

CM

CL=.1527;M=0.15

CL=0.4963;M=0.15

CL=0.1110;M=0.20

CL=0.4547;M=0.20

)LJXUH3RZHUHG/RZ6SHHG&RQILJXUDWLRQ(OHYRQ(IIHFWLYHQHVV
53

7.5 Thrust Drag Bookkeeping

In order to predict the drag from a powered solution the thrust had to be removed, 

this led to an exercise in thrust-drag bookkeeping. 

The first step was to evaluate the isolated nacelle. A closed body nacelle was 

created to eliminate bookkeeping problems that would result from tracking the mass 

flow through the nacelle. Powered and unpowered Overflow solutions were 

obtained using this isolated nacelle. The solutions are shown in Figure 51 and 

Figure 52. The powered nacelle has negative drag; it produces thrust. To obtain the 

proper drag for the configuration at low speed the nacelle drag change due to power 

had to be removed. The nacelle drag change due to power was considered part of 

the thrust. The following equations were used to develop the configuration drag 

when powered.

CD = CD* - 'CD)NAC

Where

'CD)NAC = {CD)CENTER + CD)OUTBOARD}POWER-ON - {CD)CENTER + CD)OUTBOARD}POWER-OFF

CD* Is the drag of the complete configuration from Overflow

Thirty Overflow solutions were evaluated to develop powered and unpowered drag 

polars for the OREIO configuration at low speed. Solutions were obtained for 5 

angles-of-attack with 3 elevator angles both powered and unpowered. The results 

are shown in Figure 53. There is a drag penalty that is close to 50 counts at low lift 

coefficients, decreasing to about 25 counts at high lift coefficients. The takeoff climb 

out condition is at the higher lift coefficients where the penalty is near 25 counts. 
54

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE @ LOW SPEED

M=0.20; POWER-OFF

CD=1.5 cts.

)LJXUH,VRODWHG1DFHOOH6ROXWLRQ8QSRZHUHG

Boeing Research and Technology / Aero-Sciences

ISOLATED NACELLE @ LOW SPEED

M=0.20; POWER-ON (ACTUATOR DISK)

CD=-13.1 cts.

)LJXUH,VRODWHG1DFHOOH6ROXWLRQ3RZHUHG
55

70

Boeing Research and Technology / Aero-Sciences

LOW SPEED TRIMMED DRAG POLAR

Trimmed Drag Polars

M=0.20

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

0.000 0.005 0.010 0.015 0.020 0.025

CD

CL^2

POWER-ON

POWER-OFF

)LJXUH3RZHUHGDQG8QSRZHUHG7ULPPHG'UDJ3RODUV
56

The same Overflow solutions that were used to create the low speed trimmed drag 

polars were also used to determine the elevator angle required for trim. The 

Overflow runs were completed with deflections of the center elevon only. The 

results are shown in Figure 54. For takeoff with high power the elevon would have to 

be at a negative angle of about 14 degrees. The exact angle varies somewhat with 

lift coefficient. Note that a negative angle is trailing edge up. The power-off 

solutions require a positive or trailing edge down elevon angle of about 3 degrees for 

trimming. Landing cases have so little thrust that they are very close to the poweroff cases. An estimate of the pitching moment due to thrust was included in the 

analysis to get the elevator angle required for trim.

71

Boeing Research and Technology / Aero-Sciences

LOW SPEED TRIMMED DRAG POLAR

ELEVATOR DEFL. TO TRIM

M=0.20

-20.000

-15.000

-10.000

-5.000

0.000

5.000

0.000 0.100 0.200 0.300 0.400

CL

Ge

POWER-ON

POWER-OFF

)LJXUH(OHYDWRU5HTXLUHGIRU7ULP

8 Rotor Flow Field 

The rotor flow field was evaluated for the cruise and low speed conditions. The rotor 

flow field is very important to the rotor and rotor manufacturer. The rotor flow field 

can have an effect on the performance of the rotor, the blade stresses, and the noise 

produced by the rotors. The goal would be to have a uniform flow field that is at or 

below the free stream velocity of the aircraft. The flow field for the OREIO 

configuration is not perfect.
57

The configuration used for the flow field evaluation includes the wing, body, nacelles, 

and vertical tails. The vertical tails have been twisted for near zero lift at cruise. The 

nacelles are positioned so the spacing between the rotor tips and the body upper 

surface is 3 ft and are oriented so that the rotors are normal to the cruise onset flow. 

The flow fields are shown in Figure 55 through Figure 58, for the cruise case, and 

Figure 59 through Figure 62, for the low speed takeoff condition. Four figures are 

shown for each condition. The components of velocity were converted to polar 

coordinates so that they would be more relevant to the rotor. The components are 

axial, radial, and circumferential velocity. The axial velocity is aligned with the 

nacelle axis. The radial velocity is inward or outward from the center of the rotor 

disk. The circumferential velocity is also referred to as V-theta.

The results show that the flow field is not ideal for the rotors. The circumferential 

velocity is sometimes positive and sometimes negative. This indicates that the flow 

field is sometimes clockwise and sometimes counterclockwise around the nacelle. 

The axial velocity was converted to Mach number for easier comparison to the free 

stream condition. Much of the axial Mach number ranges from about 0.83 to 0.89 

and the peak speeds are higher. The radial velocity is not shown. The axial and 

circumferential velocity and were combined to evaluate the flow angle in the front 

rotor plane. The “alpha” plotted is not the angle-of-attack that a blade would see, 

that would depend on the rpm of the rotor and the blade twist. The alpha shown is 

the flow angle that the front rotor will be immersed in. The front rotor will experience 

the alpha variation in Figure 57. Alpha varies from -2.8 degrees to +2.2 degrees. 

The last flow field figure shows the total pressure in the forward rotor plane. The 

pressure is all near free stream. There is no total pressure loss from strong shocks, 

vortices, or boundary layer ingestion. 

The flow field was improved by orienting the rotor disks normal to the onset flow. 

The alpha variation in the front rotor disk when the nacelles were aligned parallel to 

the body axis was from about -9 to +9 degrees.

The flow field of the front rotor was also checked at the takeoff climb out condition. 

An Overflow solution at Mach 0.20 and 8.0 degrees angle-of-attack was evaluated.

The results are shown in Figure 59 through Figure 62. The alpha variation in the 

front rotor disk is greater because the nacelles were aligned for the cruise onset 

flow. There is a misalignment at this condition. The alpha varies from about -5

degrees to +4 degrees. Mach number is much higher than free stream because of 

the large power setting and the large pressure jump through the rotors.

The flow field in the front rotor plane can be improved. The evaluation of the flow 

field was done late in the study and nothing was done to improve what was found. 

The nacelles were aligned normal to the power-off flow field. The power-on flow

field can be checked to see if there is any misalignment. It is possible that slight 

misalignment in the right direction may even help. The peak velocity in the rotor 

disks can be reduced by reducing the diameter of the nacelles at the rotor hub. 

Depressions in the body surface may also help. Changes in the camber of the aft 
58

body or position of the vertical tails may also help to reduce the velocity in the rotor 

disks. The development of design techniques to improve the flow field in the front 

rotor plane will have to wait for a future project.

Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES 

EFFECT OF NACELLE ALIGNMENT

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59

Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES 

EFFECT OF NACELLE ALIGNMENT

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Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES 

EFFECT OF NACELLE ALIGNMENT

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60

Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES 

EFFECT OF NACELLE ALIGNMENT

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Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES

AT LOW SPEED

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61

Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES

AT LOW SPEED

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Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES

AT LOW SPEED

)LJXUH)ORZ$QJOHLQWKH)URQW5RWRU'LVNDW7DNHRII&OLPE2XW
62

Boeing Research and Technology / Aero-Sciences

PROP FLOW ANGLES

AT LOW SPEED

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9 Concept Design Data Summary 

9.1 Final Configuration

During the initial performance evaluation of the geometry shown in Figure 7, a quick 

balance analysis showed the configuration was not viable. A chord extension was 

added to allow room for a possible aft body depression to reduce the drag of the 

nacelle installation (Section 5.1). Then when the PAI was started it was found that 

the engines needed to be located as close to the trailing edge as possible because 

the local velocity above the body continues to decrease up to the trailing edge 

(Figure 17). After the vertical tails were evaluated they were also moved aft to 

provide better acoustic shielding (Figure 29). The aft engine and tail also required 

an increased amount of aft body structure for attachment. The impact of this was to 

move the center of gravity of the configuration too far aft. To improve balance and 

loadability the aft most portion of the configuration was shortened by 80”. A new 

geometry was created that effectively removed an 80” plug forward of the engines 

but aft of the wing, and allowed workable vehicle balance and loadability. This new 

geometry is shown in the final 3-View in Figure 63.
63

)LJXUH)LQDO9LHZ'UDZLQJ

9.2 Drag Build-up

The cruise drag was built up using a standard bookkeeping system within the 

CASES program. The CASES program is a standard program used at Boeing for 

estimating aero, weight, and propulsion characteristics. It can also size a 

configuration and run mission performance. The bookkeeping system is shown in 

Figure 64. The high speed drag consists of parasite, induced, compressibility and 

trim drag. The high speed drag build up for the final configuration of the OREIO 

project is shown in Figure 65. The drag build up includes an estimated drag 

reduction for HLFC (Hybrid Laminar Flow Control) and riblets. The assumptions that 

were used for estimating the drag reduction for the application of HLFC and riblets 

are shown in Figure 66 and Figure 67. The drag build-up also includes a penalty of

4.5 counts for the open rotor nacelle and pylon as discussed in Section 7.
64

Boeing Research and Technology / Aero-Sciences

CASES Standard Buildup : CD = CDp + CDi + CDc + CDtrim

parasite induced

trim

compressibility

Cruise Drag Bookkeeping System

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13

Boeing Research and Technology / Aero-Sciences

BOEING is a trademark of Boeing Management Comp any.

Copyright © 2011 Boeing. All rights reserved. BOEING

High Speed Drag Build-Up

PROJECT NAME rnasaera0x18

Shortened OREIO

SREF (FT**2) 8010.000

FN (LBS) 37000.0

AR 5.646

SWEEP (DEG) 32.119

T/C-AVE 0.1178

S-VERT (FT**2) 442.513

F BUILD-UP (FT**2)

WING 44.5750

VERTICAL 2.5460

N&P 4.6120

HLFC & RIBLETS -5.7300

EXCRESCENCE 2.9810 (0.0648)

WING TWIST 1.1215

NACELLE INTERFERENCE 3.6050

F-TOTAL (FT**2) 53.7105

E-VISC 0.8823

CRUISE CD BUILD-UP

M-CRUISE 0.8000

CL-CRUISE 0.2650

CD0 0.00671 (0.595)

CDI 0.00449 (0.398)

CDC 0.00007 (0.006)

CDTRIM 0.00000 (0.000)

CDTOT 0.01126

L/D 23.5287

ML/D 18.8230

)LJXUH'HWDLOHG+LJK6SHHG'UDJ%XLOG8S
65

17

Boeing Research and Technology / Aero-Sciences

BOEING is a trademark of Boeing Management Comp any. Copyright © 2011 Boeing. All rights reserved. BOEING

HLFC Assumptions

WING

ͻ Applies only to outboard wing upp er surf ace.

ͻ Suction to front spar1 , th en Rn=17 million2 f rom th e front spar unless th ere is a shock b efore 

Rn=17 million. (front spar to Rn=6 million at inbo ard extent of laminar run)

ͻ Use leading edge break plus an 8 d eg. turbulent wedge as the start of span for LFC .

ͻ Assumes that th ere will be a Kru eger rath er than a slat.

ͻ Apply the scale factor f rom EET-100 HLFC study.

TWIN V ERTICAL

ͻ 40% chord for vertic al tails.

NLF NACELLE

• 25% of nacelle cowl length.

• Only use 80% of laminar area due to py lon interf erence eff ects.

NLF PYLON

• 40% of py lon chord.

• 8 deg. Turbulent wedge f rom py lon/body intersection.

• Only use 80% of laminar area due to nacelle interf erence eff ects.

70% factor applied to laminar improvements only

1 Passive system

2 Rn=17 million is a goal

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Boeing Research and Technology / Aero-Sciences

Riblet Assumptions

APPLICATION

• Apply riblets aft of transition on the wing surfaces.

• Apply riblets aft of the transition of the vertical tails.

• Apply riblets aft of the transition of the nacelles.

• Apply riblets aft of the transition of the pylons.

• Use riblets on entire centerbody

6% component skin friction improvement for riblets1

10% improvement on excrescence trend

1 BTWT riblet measurements

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66

9.3 Weight and Balance

The Group Weight Statement for the final OREIO configuration is shown as Table 8. 

The weights include the assumption that primary structure consists of PRSEUS 

(Pultruded Rod Stitched Efficient Unitized Structure) type construction.

7DEOH*URXS:HLJKW6WDWHPHQW

A quick analysis shows that the vehicle’s center of gravity, building up from Weight 

Empty through Operating Weight and Max Takeoff Weight, match reasonably well 

with the centers of pressure ranges discussed in the Wing-Body development 

section and will enable workable balance and loadability. 

9.4 Propulsion Characteristics

The purpose for this project is to provide the hybrid wing body air vehicle for flyover 

acoustic analyses, particularly for use in making flyover and community noise 

assessments using the noise data taken in the Boeing LSAF with an open rotor 

model on a hybrid wing body configuration. To have a one-to-one assessment 

basis, since the LSAF tests used the GE 36 F7/A7 model rotors, the propulsion 

performance and weights used this same rotor configuration. An OREIO propulsion 

WING-BODY 118,387

OUTERWING 61,143

CENTERBODY 51,600

AFTERBODY 5,645

WINGLET/VERTICAL TAILS 2,394

LANDING GEAR 19,500

ENGINE & NACELLE 42,923

ENGINE SYSTEMS 399

ENGINE PYLON 6,932

FUEL SYSTEM 9,209

APU 1,000

FLIGHT CONTROLS & HYDRAULICS 11,798

ELECTRICAL 3,953

PNEUMATICS & AIR CONDITIONING 7,775

ANTI-ICING 526

FURNISHINGS AND EQUIPMENT 5,498

INSTRUMENTS 1,155

AVIONICS 3,434

LOAD & HANDLING 11,664

WEIGHT EMPTY 246,548

STANDARD & OPERATIONAL ITEMS 2,242

OPERATING WEIGHT 248,790

MISSION FUEL 127,010

PAYLOAD 100,000

MAXIMUM TAKEOFF WEIGHT 475,800

OREIO

FUNCTIONAL GROUP WEIGHT (LB)
67

system data requirement is also for the data base and results to be non-proprietary. 

NASA GRC therefore created the engine performance, dimensions, and weight 

using NPSS and rotor characteristics for the F7/A7. These data are based on the 

convention where propulsion performance is based on freestream to freestream flow 

such that the forebody and afterbody surfaces scrubbed by the inflow and outflow 

from the propulsor are included in net thrust. 

9.5 Performance

The aerodynamic, weight, and propulsion data were combined to produce the 

aircraft performance. The constraints used to size the aircraft are listed in Table 9.

The constraints are consistent with the final requirements discussed in Section 10

with the exception of initial cruise altitude. In order to achieve the required initial 

cruise altitude the engine would have to be increased above the reference thrust of 

the engine supplied by NASA. NASA requested for the final performance to be 

completed without scaling the engine to make it easier to perform the acoustic 

analysis. Therefore two sets of data were generated, one with the engine scaled up 

to 41,500 lbs SLST and a second with an unscaled engine with 37,250 lbs of SLST. 

A summary of the OREIO performance with the two sizes of engines is shown in 

Table 10. The larger engine resulted with lower block fuel burn because of a better 

lift to drag ratio and lower SFC as shown in Figure 68 and Figure 69.

7DEOH25(,23HUIRUPDQFH6L]LQJ&RQVWUDLQWV

OREIO Performance Sizing Constraints

Payload: 100,000 lbs

Range: 6500 nm

Critical Field Length: 10,500 ft at SL, standard day 

Engine out climb gradient: 2.7% (3-engine aircraft)

Approach Speed: 155 KCAS

Landing field length: 5200 ft at SL, standard day

Cruise Mach: 0.80

Initial cruise altitude: 35,000 ft (achieved 32,945 ft)

Measure of Merit: maximize ton-nm/lb fuel 

At fixed range and payload this is minimize block fuel
68

Engineering, Operations & Technology | Boeing Research & Technology ERA N+2 Concept Study

ORIEO with GRC-10 Engine

110000

120000

130000

140000

150000

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

Block Fuel (lbs)

Thrust Trade for Minimum Fuel

GRC-10 

Ref Thrust

6500 nm range

100,000 lb payload

Mach 0.80

SREF = 8010 sq ft

6500 nm range

100,000 lb payload

Mach 0.80

SREF = 8010 sq ft

35 kft

ICAC

30000

32000

34000

36000

38000

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

SELECTED

Initial Cruise Altitude (ft)

POINT

37,525 lbs

New selected point allows acoustic 

analysis without scaling the engine

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Engineering, Operations & Technology | Boeing Research & Technology ERA N+2 Concept Study

200000

220000

240000

260000

280000

300000

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

Operating Empty Weight (lbs)

0.2

0.22

0.24

0.26

0.28

0.3

0.32

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

Initial Cruise CL

0.43

0.435

0.44

0.445

0.45

0.455

0.46

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

SFC

Thrust Trade (continued)

22.2

22.4

22.6

22.8

23

23.2

23.4

30000 40000 50000 60000 70000

Reference SLS Installed Thrust (lbs)

L/D

)LJXUH5HVXOWVRI7KUXVW7UDGH
69

7DEOH3HUIRUPDQFH6XPPDU\

10 Conclusions and Recommendations

A BWB configuration with an open rotor installation was developed for NASA to 

study the noise shielding features of this configuration. It is thought that noise 

shielding features of a BWB are uniquely suited to the use of an open rotor engine. 

The configuration meets the requirements established early in the project. The 

wing-body configuration was developed for a cruise Mach number of 0.80, which is 

lower than many other BWB configurations because of the open rotor engine. PAI 

was studied; some of the configuration changes were based on the PAI work. The 

PAI work also highlighted the need for future studies to determine the viability of 

open rotors on the BWB platform.

OREIO OREIO

Reference Required BWB-short BWB-short

Perf. Mcr=0.80 Mcr=0.80

GRC-10 BL GRC-10 Scaled

MTOGW 470138 475800

OEW 242517 248795

Payload 100,000 100000 100000

Fuel at Max PL 127621 127005

Range (nm) 6500 6500 6500

Block Fuel Burned 115187 114385

Ton-nm/lbf 2.82 2.84

Reference Wing Area (sqft) 8010 8010

Aspect Ratio 5.646 5.646

Span (ft) 212.7 212.7

TO Ref FN @ SL, 0.0, +27F (lbf/eng) 37524.7 41500

Number of Engines 3 3

T/W 0.24 0.26

W/S 59 59

Optimum Altitude (ft) 36257 36085

Initial Cruise Thrust Ceiling (ft) 32945 35012

Initial Cruise Buffet Ceiling (ft) 40470 40233

Initial Cruise Altitude Capability (ft) 32945 35012

Cruise type Climb-Cruise Climb-Cruise

Initial cruise altitude (ft) 35000 32945 35012

Initial cruise thrust (all engines) 21184 21058

Initial cruise mach 0.8 0.8

Initial cruise L/D 22.849 23.274

Initial cruise SFC 0.4395 0.4347

Initial cruise tau 0.956 0.955

Initial cruise CL 0.2344 0.2614

Balanced field length @ SL, 59F (ft) 10500 8118 7407

2nd Segment Gradient (%) 2.7 2.7 3.8

Landing field length (ft) 5200 6120 6212

Approach speed (KCAS) 155 141 142
70

The engine positioning was all based on the PAI work. The engine was moved aft 

because the local Mach numbers were lower near the trailing edge. The engines 

were moved up because the upper surface of the body separated when the large 

thrust levels required for takeoff were used. The engines are aligned to orient the 

rotor disk normal to the cruise onset flow. The vertical tails were moved aft to 

provide better noise shield for sideline noise from the open rotors. 

Mission performance was estimated for the final configuration. Engine performance 

was supplied by NASA GRC. The estimated weight and aero characteristics were 

supplied by Boeing. Some aerodynamic interference penalties were included for the 

installation of the open rotor engine. The flight conditions for potential noise 

measurements at takeoff and landing were included.

This study highlighted some of the problems that will be encountered with an open 

rotor engine installation. Several solutions were tried but there is a considerable 

amount of work remaining. The following is a list of areas that need more 

investigation. 

A puller or tractor engine concept may be better for the configuration. It has

advantages over the pusher concept. The pylon wake would not cause distortion 

through the rotors. Filling in the wake with a blowing system for a pusher engine 

would just add complexity. The engine airflow would also be simpler; there is added 

complexity for a push engine to get the airflow past the rotors, either externally or 

internally. A future study should include a revised rotor configuration with the right 

diameter, blade count and thrust loading. 

The rotor flow is higher than free stream and not uniform. Many aspects of the 

configuration may have to be changed to improve the flow field. The camber of the 

entire aft fuselage, the local shape of the fuselage under the nacelle and rotors, and 

the vertical tail position and camber may all have to be changed

The engine support or pylon needs more investigation. Based on the work of this 

study, it will be difficult to develop a low drag pylon installation and if the engine is a 

pusher there is the problem of wake ingestion that needs to be addressed.

The low speed upper surface separation resulted with adverse configuration 

changes. If the separation can be controlled better the engine height can be 

reduced which will improve the weight and controllability of the configuration.

The nacelle shape should be investigated. The shape used for this study was not 

designed for the cruise speed of this configuration. The nacelle shape was designed 

to provide the proper rotor flow field at Mach 0.72 for an isolated nacelle. Some 

investigation is required to determine the best shape for an installed nacelle at the 

higher cruise speed of this configuration. 
71



11 References 

1. http://lemanusa.com/wp-content/uploads/2010/01/Air-Cargo-ULD2.pdf
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Contractor Report

 4. TITLE AND SUBTITLE

Blended Wing Body Concept Development with Open Rotor Engine 

Integration 

5a. CONTRACT NUMBER

NNL10AA05B

 6. AUTHOR(S)

Pitera, David M.; DeHaan, Mark; Brown, Derrell; Kawai, Ronald T.; 

Hollowell, Steve; Camacho, Peter; Bruns, David; Rawden, Blaine K.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

 NASA Langley Research Center The Boeing Company 

Hampton, Virginia 23681-2199 Boeing Research and Technology

 5301 Bolsa Avenue

 Huntington Beach, CA 92647 

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

Langley Technical Monitor: Craig L. Nickol

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 05

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

The purpose of this study is to perform a systems analysis of a Blended Wing Body (BWB) open rotor concept at the conceptual design level. This concept will 

be utilized to estimate overall noise and fuel burn performance, leveraging recent test data. This study will also investigate the challenge of propulsion airframe 

installation of an open rotor engine on a BWB configuration. Open rotor engines have unique problems relative to turbofans. The rotors are open, exposed to 

flow conditions outside of the engine. The flow field that the rotors are immersed in may be higher than the free stream flow and it may not be uniform, both of 

these characteristics could increase noise and decrease performance. The rotors sometimes cause changes in the flow conditions imposed on aircraft surfaces. 

At high power conditions such as takeoff and climb out, the stream tube of air that goes through the rotors contracts rapidly causing the boundary layer on the 

body upper surface to go through an adverse pressure gradient which could result with separated airflow. The BWB / Open Rotor configuration must be 

designed to mitigate these problems. 

15. SUBJECT TERMS

Advanced concept; Blended wing body; Emissions; Fuel burn; Noise; Open rotor

18. NUMBER

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 PAGES

76

19b. TELEPHONE NUMBER (Include area code)

(443) 757-5802

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

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3. DATES COVERED (From - To)

September 2010 to August 2011

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

NNL10AC78T

5f. WORK UNIT NUMBER

699959.02.10.07.01 

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/CR-2011-217303

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