ac82807ec774081e6f5a6867b9344e51.ppt

- Количество слайдов: 37

Low Weight Rotor Blade Structural Design Ed Smith Jianhua Zhang Professor Research Associate Rotorcraft Center of Excellence Department of Aerospace Engineering The Pennsylvania State University June 2005

Background • A low weight rotor system is an important goal for helicopters and tiltrotors, and is an enabling technology for a cost effective large transport rotorcraft • Primary operating cost drivers are weight and power Rotor system weight: blade, hub and controls Power: low disk loading and low aircraft drag • Reduced weight and lower disk loading lead to: Larger, lighter rotors with novel hub and control concepts Radically altered dynamic characteristics • Computations by government and university personnel, and industry provide design experience needed to critique and guide work

Background Government Stability, loads, vibration, deflection, performance; compare with design criteria Industry Identify new materials and other technologies; examine feasibility and develop design guidelines University Determine minimum weight structural design for specific loads; define blade structural and inertia properties

Background Specific technologies explored for low weight rotor design • All graphite blades (compared to graphite/glass hybrid blades) • Possible Flexible composites for hubs and blades • Bearingless hub (unproven to date for large rotor system) • Composite Tailored blades & hubs for stability augmentation • Reduced or eliminated leading edge weight system (aft. cg stabilization) • Active controls for primary flight control and loads management • Additional technologies, as needed, will be evaluated and discussed by entire team

Design Parameters for Composite Blade Some design parameters which will be evaluated • Ply orientation • Ply thickness & number of layers • Spar location & designs • Leading edge mass • New materials • Active control concept

Composite Blade Cross Section Modeling Spar Web Skin Foam Composite Blade Cross Section State of Art of Modeling • Simplified composite model • Vlasov Theory Smith, Chopra Chandra, Chopra; Smith, et. al For design and beam aeroelasticity analysis • VABS (Variational Asymptotical Beam Sectional Analysis ) Hodges, et. al • 3 D shell element Analyze and detail designs

Composite Blade Cross Section Modeling • Current composite cross section model is based on Vlasov theory Translate the two dimensional plate equations into one dimensional beam equations that are only a function of the axial coordinate Originally derived for isotropic sections, and was later extended to composites Provide a method to determine the beam properties from the laminate properties and the cross section geometry • Composite modeling is based on classical laminated plate theory, where each individual lamina are summed together to form properties of the laminate. The stiffness of the individual lamina depend on their orientation within the laminate. • Blade cross section stiffness matrix and force vector terms are derived using Hamilton’s Principle, based on Vlasov beam equations

Composite Blade Cross Section Modeling – Laminated Plate Theory Classical laminated plate theory is a method to determine the properties of a laminate composed of individual lamina stacked together. The stiffness coefficients are a function of the ply angle θp and the stiffness properties of the ply itself n General Plate Segment s z The properties of a composite laminate are calculated by integrating through the thickness of the plate. The Classical relationship between the stress resultants and the linear laminate strains is given by Plate stress resultants Plate strain Plate moment resultants Plate bending curvature

Composite Blade Cross Section Modeling – Vlasov Theory for Closed Sections • Using the classical laminated plate theory as a basis, Vlasov theory is used to relate beam displacements and rotations to the beam forces to find the blade stiffnesses • The plate forces are related to the blade forces through the principle of virtual work • The plate strains ( ) and first order curvatures ( ) in the above equation can be expressed in terms of the blade displacements and rotation through geometric considerations • The generalized blade force to generalized blade displacements relation is derived as : Axial force (Extension stiffness) : Lag bending moment : Flap bending moment (Flap bending stiffness) : Torque (Lag bending stiffness) : Bimoment (Torsion stiffness)

Design Procedure • Identify design variables – selecting materials, skin and spar thickness (No. of plies), web location, ply orientation at 4 radial locations (r/R= 0. 25, 0. 50, 0. 75, 1. 00) • Thickness of the skin and spar are changed by increasing or reducing number of plies. The ply angle for the spar starts at 0 and the ply angle for the skin starts at ± 45. The ply angles of skin and spar will be varied to meet the stiffness and strength requirements • The design process will continue until the blade cross section inertia and stiffness properties are within the targeted range, and the stresses or strains satisfy the failure criteria • Non structural mass is assumed to be zero; it will be added in the comprehensive analysis of the rotor system Spar Web Skin Foam 1. 0 0. 75 0. 5 r/R 0. 25 Composite Blade Cross Section

Design Procedure Materials; airfoil; cross section configuration; sectional loads Design variables: skin, spar thickness, ply angles, web location Adding or reducing # of plies of skin and spar and changing ply angles Analysis of cross section stiffness and stress via thin walled composite beam theory Targeted blade cross section properties: Inertia, stiffness, etc cg offset and amount of non structural mass needed Satisfied ? Strength or Strain criteria no yes Output blade cross section inertia, stiffness, offsets, etc. for the next design iteration

Blade Cross Section Loads ● The blade loads supplied by NASA are based on the speed sweep (up to the total rotor power 15000 hp) and the load factor sweep (up to 1. 54 g). Blade Loads at different flight conditions will be evaluated and the worse case will be used in the blade design ● All six section load components: flap bending moment, lag bending moment, torsion moment, chordwise force, normal force and axial force are given at designated radial locations in the four forms – max. , min. , average, 1/2 Peak ● Based on these loads provided, the worst loading conditions are sorted out by assuming all six components reach the largest at the same time ● In the current stress or strain analysis, the normal shear loads and chordwise shear loads are neglected

Blade Cross Section Stress Analysis Stress/Strain calculation is based on the worst loading cases and are calculated at the middle of each layer of each segment along the cross section ● The stiffness matrix is derived from the composite blade modeling, and the blade cross section loads are from the rotor comprehensive analysis. Then, the displacements can be found using a linear solver ● Once the blade displacements are known, the blade strains and curvatures can be found by geometric consideration ● Finally, using laminated plate theory, the stress distribution across the blade cross section can be obtained Composite Blade Cross Section Modeling Blade Loads Blade Displacement Geometric Consideration Plate Strain & Curvatures Laminated Plate Theory Stress Distribution Across Blade Cross Section

Design Criteria Strength • Macromechanical failure theory ▪ Maximum stress ▪ Tsai Hill (Deviatoric energy theory) ▪ Maximum strain ▪ Tsai Wu (Interactive tensor theory) • Tsai-Wu strength failure criterion is applied in the strength analysis normal stress in longitudinal direction normal stress in transverse direction shear stress X longitudinal tensile strength X’ longitudinal compressive strength Y transverse tensile strength Y’ transverse compressive strength Limitations: Does Not Address Laminate Failure Modes - Delamination - Damage Tolerance (Holes, Notches, etc. )

Design Criteria – Laminate Strain Allowables Industrial Design Practice • Typically, an allowable of 3000 microstrain is a laminate allowable associated with a particular lay up (usually quasi isotropic) with all kinds of knock downs. The lamina level strength values are not typically referred to as allowable and not used in design • Carbon fiber design strains for aircraft structure are typically in the range of 3000 4500 microstrain range because that has been found to provide a realistic conservative design allowable for a damaged structural laminate under cyclic loading • Current allowables for IM 7/8552 are on the order of 4500 microstrain (compression) and 6000 microstrain (tension). However, the factor of safety (or some may say “ignorance”) reduces these to less than 3000 microstrain in design Industry design practice of 3000 microstrain allowable will be adopted for the current blade structural design

Blade Structural Design • A composite rotor blade modeling program has been developed and adapted for blade cross section design. Detailed cross section stress/strain analysis, have been formulated and applied in the design process. Design criteria have been established • The baseline blade properties were from the scaled XV 15 blade, from which the initial blade loads were calculated; based on the loads and other design requirements, such as stiffness, C. G. location, etc. , a new blade design will be conducted and the blade properties will be fed back for comprehensive analysis until the design process converges • Ten design iterations of LCTR blade have been accomplished; the newly designed blade properties satisfy the requirements from rotor comprehensive analysis, and the overall weight reaches the targeted 50% reduction • Two iterative design process for LABC and one for LCTC have also been completed

Composite Materials Properties • For the preliminary studies, AS 4/3501 composite materials were used; then a new advanced composite material was postulated with high modulus and high strength (1. 67 times higher than AS 4/3501 6) to study the influence of material properties on low weight blade design • IM 7/8552 is chosen in the final design because of its higher modulus and higher allowables AS 4/3501 -6 (with factor of 1. 67) IM 7/8552 Unit 20. 6 34. 4 23. 8 msi 1. 49 2. 49 1. 7 msi 1. 04 1. 74 0. 754 msi 0. 27 0. 32 1. 7718 10 3 Slugs/cubic inch Xt 331 553 395 ksi Xc 208. 9 349 245 ksi Yt 8. 3 16. 6 16. 1 ksi compared to AS 4 Yc 33. 1 99. 3 21. 8* ksi S 10. 3 17. 2 17. 4 ksi Density Torsion stiffness Requirement is hard to meet Strength Almost doubled * Generic IM 6/Expoxy UD prepreg Sources: 1. Engineering Mechanics of Composite Materials. Isaac M. Daniel and Ori Ishai, Oxford University Press, 1994. 2. Hexply 8552 from Hexcel Composites 3. Industries and Government

Blade Cross Section Design LCTR 0. 75 r/R: 0. 25 0. 75 1. 0 t/c: 0. 20 0. 18 0. 12 0. 08 • Blade with different thickness ratios and taper ratios were investigated for their effects on the blade weight. Tip to root taper ratio of 0. 8 was chosen for LCTR • The blade was designed at four radial locations (r/R=0. 25, 0. 75, 1. 0)

Blade Cross Section Design LCTR Stiffness requirements 1. 0 Blade Loads 0. 75 Design drivers 0. 5 r/R 0. 25 0. 75 1. 0 Design loads level 60 loads turn 80 sls rotor 2 • Design drivers are different at each cross section. In the current design, the design drivers are either load limits or the stiffness requirements • The worst loading condition is chosen for each cross section by evaluating the max. loads among different flight conditions

Blade Cross Section Design LCTR Cross section (the 10 th iteration) Radial station (r/R) Chord (c) (ft) Thickness ratio (t/c) Skin (trailing edge) Skin+Spar Web location (%c) 0. 25 3. 5 0. 20 [(± 45)10] [(± 45)33/(0)60/(± 45)2] [(± 45)35] 40 0. 50 3. 3 0. 18 [(± 45)10] [(± 45)33/(0)30/(± 45)2] [(± 45)35] 40 0. 75 3. 1 0. 12 [(± 45)5] [(± 45)28/(0)20/(± 45)2] [(± 45)30] 40 1. 00 2. 9 0. 08 [(± 45)5] [(± 45)28/(0)10/(± 45)2] [(± 45)30] Ply thickness (in) 40 0. 005 • Uniform skin lay ups for 0 50%R and 50 100%R for trailing edge • Slight skin taper for leading edge • Moderate spar taper

Cross Section Strain Analysis Normal Strain (Microstrain) r/R=0. 25 root Min ( 1800) Max (2800) r/R=0. 50 mid span Min ( 1300) r/R=1. 0 tip r/R=0. 75 Max (2900) Min ( 2300) Max (2200) Max (2900) Min ( 2000) The normal strains across the section are all within 3000 micro strain

Cross Section Stress Analysis Normal stress (Ib/sq. foot) r/R=0. 50 mid span r/R=0. 25 root Max (9. 2 E+6) Min ( 5. 9 E+6) Max (6. 9 E+6) r/R=1. 0 tip r/R=0. 75 Max (9. 2 E+6) Min ( 6. 7 E+6) Min ( 4. 2 E+6) Max (8. 7 E+6) Min ( 4. 9 E+6) Tapered airfoil thickness Stresses shown are calculated based on the loading condition when the blade cross section loads: axial force, flap and lag bending moments, and torsion moment are all the largest

Blade Cross Section Properties • Moderate taper of spar and tapering of blade chord and thickness save blade weight, but the stiffnesses also drop quickly, especially the blade flap stiffness • The blade has thick torque box to meet the requirement of high torsion stiffness

Sensitivity Studies – Blade Load Reduction • Recent studies show that blade loads can be significantly reduced by the active blade management, such as the concept of dual active trailing edge flaps • In the current sensitivity study, the flap bending moment is reduced by 50% Deformed blade w/o control Opposite action of dual flap lift due to outboard flap Opposite lift due to inboard flap Straightened blade Dual flap concepts − Generate additional moments Results in reducing blade load Reduce blade stresses and increase blade life − Effect to trim by dual flap could be minimized (net lift is nearly zero) Control inputs include 1/rev and higher harmonic components

Sensitivity Studies – Blade Load Reduction r/R 0. 25 0. 50 0. 75 1. 0 Overall blade weight reductions Weight reduction 21% 15% 13% 12% 18% Reducing the flap bending moment by 50% can save 18% of weight. It is most effective at the blade root, where the flap bending moment is the largest

Sensitivity Studies Materials The new advanced composite materials with high modulus and high allowables are expected in the future, therefore some sensitivity studies of materials are essential for the success of low weight rotor design • Increase the ultimate strengths by 50% via improved materials and better damage design and detection • Increase elastic modulus in fiber direction by 25% via using stiffer fibers

Sensitivity Studies Materials r/R 0. 15 0. 55 0. 75 0. 98 Overall reduction 50% increase of ultimate strength 23% 14% 12% 10% 25% increase of El (longitudinal) 0% 16% 18% 13% 8% • Increasing the ultimate strengths by 50% achieves about 20% weight reduction; It is most effective at the blade root, where the flap bending moment is the largest, and also the driving factor for blade section design • Increase the material elastic modulus by 25% is not effective at those blade section, where driving factor is the loads (0. 15 R), but they are as effective at those blade sections, where the stiffness is the driving factor (0. 50, 0. 75 and 1. 0 R for example). The Overall weight reduction is about 8%

Blade Design Off axis Plies in Spar • The structural design iterations on the blade have included spars with 0° plies and skins with ± 45° plies. According to industries’ comments, there needs to be some off axis plies in the spar, in order for it to provide torsional resistance to the applied loads • The current LCTR blade design requires high torsion stiffness, the off axis plies in the spar will be investigated for their influences on the blade stiffness, especially how they contribute to torsion stiffness • The spar with 50% of off axis plies will be investigated for three cases with different ply angles: ± 10°, ± 30°

Blade Design Off axis Plies in Spar Torsion Stiffness Initial design Flap Stiffness 50% of ± 10° plies in spar Lag Stiffness 50% of ± 30° plies in spar • The largest torsion stiffness increase is achieved at root section about 15% for the case of 50% of ± 30° plies in spar • The flap and lag stiffness decrease according (28% and 11%)

Blade Design CG Placement via Tailoring Spar • Tailoring spar topology to move CG forward such that less dead weight needed for stability • Add more ± 45° plies at leading edge such that the CG location can be pushed forward and the blade torsion stiffness can be increased

Blade Design CG Placement via Tailoring Spar Torsion stiffness mass C. G. Location Initial Design Quarter of Chord After CG placement design • By adding more ± 45° plies at leading edge, CG locations at all four cross section are ahead of quarter of chord • Torsion stiffness has been significantly increased by 35% while the sectional mass has also been increase by 35%

Summary – Blade Design • A composite rotor blade modeling program has been developed and adapted for blade cross section design. The program is capable of calculating blade inertia and stiffness properties, all offsets, and the stress and strain distribution across the blade cross section based on the blade loads provided • After discussion with industries and government, the blade design criteria have been established. The Industry design practice of 3000 microstrain allowable is used for the current blade structural design studies • The baseline blade properties were from the scaled XV 15 blade, from which the initial blade loads were calculated; based on the loads and other design requirements, such as stiffness, C. G. location, etc. , a new blade design will be conducted and the blade properties will be fed back for comprehensive analysis until the design process converges

Summary – Blade Design • For preliminary design, AS 4/3501 composite material was used; IM 7 was used in final design for its high modulus and high strength. In order to investigate the material in the future • Blade with different thickness ratios and taper ratios were investigated for their effects on the blade weight. Tip to root taper ratio of 0. 8 was chosen for LCTR • Ten design iterations of LCTR blade have been accomplished; the newly designed blade properties satisfy the requirements from rotor comprehensive analysis, and the overall weight reaches the targeted 50% reduction • Two iterative design process for LABC and one for LCTC have also been completed

Summary – Sensitivity Studies • Sensitivity studies of material properties on blade weight – Increasing the ultimate strengths by 50% achieves about 20% weight reduction; It is most effective at the blade root, where the flap bending moment is the largest, and also the driving load for blade section design – Increase the material elastic modulus by 25% is not effective at those blade section, where driving factor is the loads (0. 15 R), but they are as effective at those blade sections, where the stiffness is the driving factor (0. 55, 0. 75 and 0. 98 R for example). The Overall weight reduction is about 8% • Study of active loads control (active trailing edge flaps) on blade weight – Reducing the flap bending moment by 50% can save 18% of weight. It is most effective at the blade root, where the flap bending moment is the largest.

Summary – Design Issues • Off–Axis Plies in Spar Using off axis plies in the spar can increase the torsion stiffness and decrease the flap and lag stiffness. This method can be used for tailoring blade frequencies • C. G. Placement by Tailoring Spar CG placement by adding more ± 45° plies at leading edge can push the CG ahead of quarter of chord and also increase the torsion stiffness, however, the sectional mass also increases. Therefore, it may help for CG placement, but it may not change the torsion frequencies

Summary – Research Topics Materials • The new advanced composite materials with high modulus and high allowables are essential for the success of low weight rotor design. It is a challenging task to predict the composite material development in the next 15 years • The current blade design uses all Graphite material. As an opposite, all glass blades could be designed to explore the design boundary; and a hybrid blade (Gr. + glass) could be investigated for the tradeoff and benefits of using multiple composite materials for blade design Blade design Details The leading edge protection cap, the trailing edge block, the blade dead weight, as well as anti icing blanket should be taken into consideration for design in order to more accurately estimate the blade weight

Summary – Research Topics Modeling Enhancement & Validation • An automated blade cross section optimization programming will be developed. It is expected that by using this optimization program, the topology of the spar and skin can be further optimized to reduce the blade weight and the design process can be more efficient • Validation studies with other more sophisticated composite blade modeling algorithms such as VABS (In progress) All researches will be documented in the final report to NASA as well as the following three conferences: • AHS October 2005 Rotorcraft Structures and Survivability Specialist Meeting • AHS November 2005 2 nd International Basic Research Conference • AHS January 2006 Design Conference