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Airframe Development for the Hybrid Wing Body Aircraft

Alex Velicki 1 and Patrick Thrash 2 The Boeing Company, Huntington Beach, California 92647-2099

Dawn Jegley 3 NASA Langley Research Center, Hampton, Virginia 23681-2199

By using the pressure cabin to generate lift, the airfoil-shaped Hybrid Wing Body (HWB) aircraft offers a revolutionary new way of improving overall airplane performance. It moves beyond traditional weight-savings schemes by radically reshaping the fuselage centerbody to achieve a higher order lift-to-drag ratio than would otherwise be possible using a conventionally shaped fuselage. Although such a reshaped centerbody is aerodynamically desirable, this transformation imposes significant weight and cost challenges onto the airframe structure. This paper describes how researchers at NASA and Boeing are working together to characterize and resolve these fundamental weight and producibility issues for the HWB airframe.

Nomenclature Isp = specific fuel consumption L/D = lift-to-drag ratio Nx = X component of the internal running load acting in the streamwise direction Ny = Y component of the internal running load acting in the spanwise direction Nz = Z component of load acting out-of-plane caused by internal pressure Wf = final aircraft landing weight Wi = initial aircraft take-off weight

I. Introduction he first 100 years of transport airframe development have largely been characterized by the creation of higher performing material systems, that when coupled with improvements in engine technology, have been used to

produce ever larger, more powerful, fuel-efficient aircraft. This trend has led to a revolution in air travel and global economic expansion far beyond anything ever imagined by those early aviation pioneers. Yet, the techniques they employed, of developing better materials and then exploiting them on the next tube-and-wing airplane design, are still the same methods in use today. Surprisingly enough, even after 100 years of innovation, this particular aspect of the design process remains unchanged. The main problem with such an approach is that meaningful airframe advancements have now become hostage to the physics-based limitations imposed by achieving true material system breakthroughs which become successively more difficult over time.

This progression is evident from the outset of those first few flights in a wooden airframe. From that start, there have only been two revolutionary changes in airframe design - first in the 30-year march to the metallic skin-stringer designs realized in the 1930’s, and then some 75 years later, in the shift to the carbon-fiber airframe of the 787 (Fig. 1). This long and lengthening development cycle illustrates the inherent problem in the materials-discovery improvement cycle - and it implies that the next fundamental shift in airframe technology may not occur for

1 Principal Design Engineer, The Boeing Company, Advanced Structures R&D, 5301 Bolsa Avenue. 2 Principal Manufacturing Engineer, The Boeing Company, Advanced Manufacturing R&D, 5301 Bolsa Avenue. 3 Senior Aerospace Engineer, NASA-LaRC, Structural Mechanics and Concepts Branch, AIAA Associate Fellow.

T

Figure 1. 100 years of structural design evolution.

47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition5 - 8 January 2009, Orlando, Florida

AIAA 2009-932

Copyright © 2009 by The Boeing Company. All rights reserved. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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another 150 years! Clearly, if the aviation community hopes to some day cap, or even maintain current emission levels while growing traffic, a far more effective means of improving airframe performance must be considered beyond the mere material-substitution approach used in the past.

Moving beyond incremental improvement will require a more far-reaching consideration of the airplane design space. It must include such revolutionary airplane designs like the HWB, or Blended Wing Body (BWB) concept which represents just the sort of performance improvement that is possible when the conventional airframe design is pushed beyond a tube-and-wing configuration (Ref. 1). Such a step function in performance improvement becomes possible because the airfoil-shaped fuselage of the HWB generates a higher L/D coefficient than is possible with a conventional circular fuselage. The relative improvement can easily be seen in the Breguet range equation because range, or improved performance, is directly proportional to the improvement in L/D as long as the remaining variables are held constant. (Fig. 2)

For a typical HWB-type planform, where roughly 20% of the overall lift is generated by the centerbody region of the design, there is a corresponding improvement in range. Using this design approach, the challenge now becomes designing a non-circular pressurized payload cabin that is not only lightweight, but also economical to produce - making the airframe design the enabling technology for achieving vehicle closure.

To meet these challenges, researchers at The Boeing Company in Huntington Beach (Phantom Works Division) and at NASA Langley Research Center (NASA-LaRC) are developing a highly-integrated stitched-composite airframe solution that is tailored and optimized for the HWB airframe by exploiting the unique processing advantages inherent in dry carbon fabrics and the damage-arrest characteristics of stitched structures (Ref. 2). This new design and manufacturing concept is called the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS). It is a disruptive technology that is a conscious departure from conventional laminated composite design practice that was developed to meet the demanding structural performance and producibility requirements of the HWB design space. This paper summaries the development approach that is being pursued to establish the fundamental structural characteristics of the PRSEUS HWB airframe.

II. HWB Airframe Design Fundamentals An essential feature of the HWB structural approach must be the capability to economically produce integrated

compound-curvature panels. The variable loft geometry makes the cost of tooling individual structural elements (skins, stringers, frames, bulkhead caps, etc.) prohibitively expensive. As such, this condition precludes the use of conventional composite fabrication techniques that require numerous detail and subassembly tools to lay-up individual parts and then subsequently bond them together into completed panel assemblies. Clearly, structural solutions with less tooling are appealing, and indeed this single aspect formed the basis on which the PRSEUS structural concept was developed.

Another key difference for the HWB airframe is the unusual bi-axial loading pattern in the shell (Fig. 3). Here the load magnitudes are more nearly equal in each direction (Nx and Ny) than what is normally found on conventional tube-and-wing fuselage arrangements when the cantilevered fuselage is more highly loaded in the Nx direction, along the stringer, than in the Ny direction, along the frame. This single difference has a profound impact on the structural concept selection because it dictates that the optimum panel geometry should have continuous load paths in both directions (Nx and Ny), in addition to efficiently transmitting internal pressure loads (Nz) for the near-flat panel geometry. For conventional skin-stringer-frame built-up panels, the frame shear clip is made discontinuous to allow the stringer

Figure 2. Centerbody used to generate lift.

Figure 3. Pressure cabin internal running loads.

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to pass through uninterrupted in the primary longitudinal loading direction. If such an arrangement were used for the HWB, then the frame member (attached by a discontinuous shear clip to the skin) would be less effective in bending and axial loading, ultimately resulting in a non-competitive solution.

To combat this problem, the HWB PRSEUS fuselage panel has been designed as a bi-directionally stiffened panel design, where the wing bending loads are carried by the frame members and the fuselage bending loads are carried by the stringers. Features of this design include: loads paths that are continuous in both directions, skin and flange laminates that are highly tailored, thin skins designed to operate in the post-buckled design regime, and stitched interfaces to arrest damage propagation. The resulting panel design is extremely effective in eliminating the weight penalty associated with the non-circular HWB pressure cabin. (Ref. 2)

III. PRSEUS Structural Solution The highly integrated nature of PRSEUS is evident in the strategic placement of the carbon fibers (Fig. 4). The

dry warp-knit fabric, pre-cured rods, and foam-core materials are assembled and then stitched together to create the optimal structural geometry for the HWB fuselage loading. Load path continuity at the stringer-frame intersection is maintained in both directions. The 0-degree fiber dominated pultruded rod increases local strength/stability of the stringer section while simultaneously shifting the neutral axis away from the skin to further enhance the overall panel bending capability. Frame elements are placed directly on the IML skin surface and are designed to take advantage of carbon fiber tailoring by placing bending and shear-conducive lay-ups where they are most effective. The stitching is used to suppress out-of-plane failure modes, which enables a higher degree of tailoring than would be possible using conventional laminated materials.

The resulting integral structure is ideal for the HWB pressure cabin because it is a highly efficient stiffened-panel geometry in three directions that is damage tolerant, stitched to react pull-off loads, and also capable of operating well into the post-buckled design regime which enables thin-gauge skin-stringer designs to be lighter than non-buckled sandwich designs. The PRSEUS HWB airframe features large unitized wing and fuselage components that offer efficient continuous load paths, higher notched design properties, and larger allowable damage levels with enhanced levels of survivability beyond those possible using unstitched designs (Fig. 5). The primary structural assemblies are PRSEUS-based designs (red region) and the non-pressurized areas (green region) are baselined as a combination of metallic and composite components.

The nexus of the PRSEUS fabrication approach (Fig. 6) is the self-supporting stitched preform assembly that can be fabricated without exacting tolerances, and then accurately net-molded in a single oven-cure operation using high-precision outer moldline (OML) tooling. Since all the materials in the stitched assembly are dry, there are no out-time, or autoclave limitations as in prepreg systems, which can restrict the size of an assembly because it must be cured within a limited processing envelope.

Resin infusion is accomplished using a soft-tooled fabrication method where the bagging film conforms to the inner moldline (IML)

Figure 4. Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS).

Figure 5. HWB structural breakdown.

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surface of the preform geometry and seals against a rigid OML tool, thus eliminating costly internal tooling that would normally be required to form net-molded details. The manufacture of multiple PRSEUS panels (Ref. 3, 4) proved that the essential feature of this concept – the self-supporting preform that eliminates interior mold tooling – was feasible for the near-flat geometry of the HWB airframe. This accomplishment represents a fundamental breakthrough in addressing the producibility needs for the HWB airframe.

IV. Airframe Structural Sizing To assess the structural efficiency of a PRSEUS airframe, a preliminary design trade study (Fig. 7) was

performed with a reference baseline HWB vehicle (Ref. 4) using an FEM-based global-local sizing approach that constrains the overall vehicle-level stiffness and internal load distributions in order to satisfy airplane-level maneuver requirements. The local panel-level constraints, such as panel stability and strength, were checked using local sizing codes. Updated panel stiffness and mass properties were then passed between the global and local analyses until a converged solution was reached. Once the analytical models converged, the final structurally-sized results were then multiplied by non-optimum factors to generate a final “as-fabricated” weight distribution for the airplane.

These results were then compared with prior sandwich panel trade studies - revealing the benefits of the PRSEUS approach. The pressurized shell elements (skin panels and frames added together) were 28% lighter for the PRSEUS concept than comparable sandwich panel designs. This weight reduction was due mainly to the advantages of the skin-stringer design with respect to damage arrestment, post-buckling of the skins, and an integral frame design that was more effective in bending.

In the upper cover regions of the centerbody, the confluence of

Figure 6. PRSEUS panel fabrication sequence.

Aeroelasticity

NASTRANSolution 200

Constant Set of Running Loads

Panel Optimization

Global AnalysisLocal Analysis

StiffnessMass Distribution

Linear Aero(NASTRAN Sol 144)

Local Panel Sizing Req'mts• Strength and Stability• Local Geometry Effects• Material and Mfg Constraints

Fuselage Design Region

Panel Geometry

Vehicle Level Requirements

Material Properties

Property Update

Panel Level Requirements

Fuselage Concept

Design Values from ACT Wing Testing

Embedded Sizing Equations

Bottoms-up Wts

Vehicle FEM

Aeroelasticity

NASTRANSolution 200

Constant Set of Running Loads

Panel Optimization

Global AnalysisLocal Analysis

StiffnessMass Distribution

Linear Aero(NASTRAN Sol 144)

Local Panel Sizing Req'mts• Strength and Stability• Local Geometry Effects• Material and Mfg Constraints

Fuselage Design Region

Panel GeometryPanel Geometry

Vehicle Level Requirements

Material Properties

Property Update

Panel Level Requirements

Fuselage Concept

Design Values from ACT Wing Testing

Embedded Sizing Equations

Bottoms-up Wts

Vehicle FEM

Figure 7. Global-local structural optimization overview.

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compressive axial loading and out-of-plane panel deflections are particularly acute (Fig. 8). Here, the non-linear panel response makes frame stability the critical discriminator within this design space. The combined loading of internal pressure (out-of-plane) and axial compression (along the frame direction) caused by the maneuver loads, produces large secondary-bending effects in the near-flat cover panels. As such, the combined loads are most efficiently reacted by tall continuous frame members. Structural concepts without an effective frame element straddling the spanwise region of the pressure cabin, as well as lacking an efficient means of transferring out-of-plane pressure loading (between the skin and frame elements), will not be as effective as an integral design solution like PRSEUS, where the skin and frame elements acts as a single continuous bending member.

This combined-loads environment of the pressure cabin represents the fundamental challenge for the HWB airframe, and as such, it forms the basis of the building-block structural development program formulated by NASA and Boeing researchers. By isolating and understanding each aspect of the combined-loading envelope, researchers will be able to combine individual element and subcomponent test results with complicated non-linear finite element solutions that will ultimately be used to predict and validate the complicated nonlinear behavior that occurs at the airplane level.

V. Airframe Development Testing Using the building-block approach, a series of experiments have been developed to verify that the stitched

specimens fabricated without the benefit of autoclave pressure behave as anticipated and meet the HWB design requirements. Coupon, single-stiffener, multi-stiffener and subcomponent specimens will be used to demonstrate the behavior of complex parts representative of HWB aircraft structures. (Fig. 9)

Figure 8. Combined loading for upper cover.

Figure 9. Building-block structural testing leads to combined loads test.

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Coupon-level specimens have verified that the strength and stiffness properties for out-of-autoclave composites are equivalent to those cured in the autoclave. Representative material properties from stitched coupons fabricated with an autoclave, as well as oven-cured specimens, can be found in Ref. 5, 6. A series of single-stiffener specimens (Fig. 10) were used to demonstrate the ability to predict pre- and post-buckling behavior and show that specimens can buckle without breaking stitches, therefore preventing delamination growth and forcing the flanges to remain attached to the skin. These types of experiments are described in Ref. 7, 8. Photographs of single stiffener specimens prior to testing are shown in Fig. 10 and post-test photographs for the single-stringer specimen are shown in Fig. 11. The photographs show that the damage is contained within a small area of the panel with little skin-stiffener separation.

Flat multi-stiffener specimens were used to examine behavior of pristine and damaged panels subjected to unidirectional loading. A key aspect of the PRSEUS design and fabrication approach is the ability of the structure to turn cracks in a manner similar to metal structures, thereby redistributing internal loads to reduce the crack-tip stress intensity. This characteristic was demonstrated on a representative fuselage test panel (Fig. 12). The 3-stringer dog-bone tension panel with a 2-bay crack (slot across central stringer) was statically tested to failure. This specimen was able to arrest propagating damage in both the horizontal and vertical directions as it emanated from the slot edge. Because damage was arrested at the stitched interfaces and fully contained within the 2-bay damage zone bordered by the adjacent stringers and frames, the panel was able to continue bearing load beyond the design limit load requirement. The specimen ultimately failed at the frame flange at a load level 30% higher than the design requirement, and 60% higher than the onset of damage propagation at the slot edge. This test clearly demonstrated that substantially higher loads are possible (60% higher) when stitched interfaces are used to turn cracks and arrest damage, thereby enabling a fail-safe design methodology to be employed for stitched composite structures. (Ref. 3)

Additional load path screening tests were also conducted to assess the out-of-plane load-carrying capability of the PRSEUS concept. In each case, when stringers, frames, or the stringer-frame intersection specimens were tested, the resulting failure loads far exceeded the design loads for the 2P internal pressure condition on the HWB aircraft. As can be seen in the series of photographs (Fig. 13), the integral stitched interfaces were able to suppress resin failure modes, and in most cases, resulted in fiber failures in the individual structural details rather than the first-ply

Figure 10. Stringer and frame test specimens.

Figure 11. Compression loading along stringer.

Figure 12. 3-stringer tension panel damage-arrest test.

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failure modes that are common in conventional unstitched composite constructions. These results validated the out-of-plane structural integrity of the PRSEUS concept and opened the way for further HWB airframe development activities in the future. Additional multi-stiffener minimum-gauge-skin-thickness panels are being fabricated under NASA contract NNL07AA48C (Ref. 4) which will be used to examine the behavior of panels loaded independently in axial compression, tension, and internal pressure.

In order to examine complex load combinations such as pressure and tension or compression, a larger representative fuselage section is being devised. Plans are being formulated to design and fabricate such a specimen. A sketch of the multi-bay box structure is shown in Fig. 14. This multi-bay box must be near full-scale to demonstrate the capability of the structure. In the sketch, the cover panels and floor structure are PRSEUS, while the ribs are anticipated to be simple sandwich panels. The test article represents a segment of the HWB centerbody. In addition to the evaluation of the combined-loads environment, corner joints and transitions will also be an important aspect of this test. Internal pressure up to 2P will induce pillowing effects and load the joints between ribs and covers. A series of experiments on this box would include loading a pristine box, inflicting discrete-source damage and failing the box in a combined-loads condition. This experiment is anticipated to be conducted in the NASA Langley Combined Loads Test Facility (COLTS) to safely introduce the required internal pressure load simultaneously with the tension, compression or torsional maneuver loads. A conceptual sketch of the proposed test article (Fig. 14) is shown loaded in the COLTS test fixture (Fig. 15). 9

VI. Conclusion The HWB design approach represents an

opportunity to move beyond incremental airplane performance improvements by radically reshaping the aircraft centerbody to achieve first-order aerodynamic improvements in L/D. The success of such an approach rests squarely on the development of an advanced structural concept that is capable of offsetting the structural weight and cost penalties inherent in the non-circular centerbody design. To meet these challenges, researchers at NASA and Boeing have developed a highly engineered structural solution that moves beyond traditional composite design practices to offer a highly efficient structural solution that can be operated beyond conventional no-growth design constraints. The result is a very

Figure 13. Out-of-plane loading of stitched structural interfaces.

Figure 14. Combined-loads specimen concept.

Figure 15. Specimen arrangement in COLTS facility.

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efficient airframe structure that combines the skin, stringers, and frame elements into a integral structural solution for reacting the complex loading of the HWB airframe, as well as addressing the producibility challenges. This unique approach represents a bold vision in composite design theory and manufacturing methods. It is a departure from conventional multi-detail laminated and bonded composite assemblies and progresses to larger one-piece cocured panel designs with seamless transitions and damage-arrest stitched interfaces. Characterization of this innovative approach to airframe shaping and structural concept development is the primary objective of the coordinated research work being performed by NASA and Boeing researchers under the structures portion of the NASA Subsonic Fixed Wing Technology initiative.

Acknowledgments The authors wish to thank the aerodynamic community, and in particular Dr. Liebeck and Dr. Fay Collier, for

their tireless efforts in promoting the BWB/HWB airplane design and challenging the structures community to develop suitable next-generation composite design and fabrication technologies under the NASA Fixed Wing Project.

References 1 Liebeck, R. H., “DESIGN OF THE BLENDED-WING-BODY SUBSONIC TRANSPORT”, 2002 Wright Brothers

Lecture, AIAA Paper 2002-0002, August 2002. 2 Velicki, A. and Thrash, P.J., “Blended Wing Body Structural Concept Development”, Royal Aeronautical Society

Aircraft Structural Design Conference Challenges for the Next Generation – Concept to Disposal, 14-16 October 2008, Liverpool, UK.

3 AIR VEHICLE TECHNOLOGY INTEGRATION PROGRAM (AVTIP), Delivery Order 0059: Multi-role Bomber

Structural Analysis, AFRL-VA-WP-TR-2006-3067, Krishna Hoffman, MAY 2006, Final Report for 14 December 2004 – 08 May 2006, AFRL-VA-WP-TR-2006-3067.

4 Velicki, A. and Thrash, P.J., “Damage Arresting Composites for Shaped Vehicles", NASA Contract NNL07AA48C,

Phase I Final Report for October 2007 through September 2008, released November 2008. 5 Hinrichs, S., “General Methods for Determining Stitched Composite Material Stiffnesses and Allowable Strengths,

Vol I,” McDonnell Douglas Technical Report MDC94K9113, Long Beach CA, March 1995. 6 Warp/knit Multi-Axial Carbon Fiber Fabric, The Boeing Company, Process Specification DMS 2436H, Long Beach

CA, 2005. 7 Velicki, A. and Thrash, P.J., “Advanced Structural Concept Development Using Stitched Composites”, 49th

AIAA/ASME/ASCE/SHS/ASC Structures, Structural Dynamics, and Materials Conference, 7-10 April 2008, Schaumburg, IL, AIAA Paper 2008-2329.

8 Jegley, D.C., Velicki, A., Hansen, D., “Structural Efficiency Of Stitched Rod-stiffened Composite Panels With Stiffener

Crippling”, 49th AIAA/ASME/ASCE/SHS/ASC Structures, Structural Dynamics, and Materials Conference, 7-10 April 2008, Schaumburg, IL.

9 Ambur, D. A., Rouse, M., Starnes, J.H., and Shuart, M. J., "Facilities for Combined Loads Testing of Aircraft Structures

to Satisfy Structural Technology Development Requirements. Presented at the 5th Annual Advanced Composites Technology Conference, Seattle, WA, August 22-36, 1994.