university of texas vsp structural analysis module update...

© 2012 Armand J. Chaput

University of Texas VSP Structural Analysis Module Update - Demonstration

Inaugural VSP Workshop, San Luis Obispo, CA

Sarah Brown Jose Galvan Tejas Kulkarni

Armand J. Chaput

Department of Aerospace Engineering and Engineering Mechanics, University of Texas at Austin

23 August 2012


(1)  Expand VSP user capabilities for employing higher order, physics based tools and methods during conceptual design (CD)

(2)  Integrate VSP with an open source finite element method (FEM) structural analysis program in a Model Center Environment - Focused on CalculiX (available under terms of GNU General Public

License as published by the Free Software Foundation) (3) Facilitate application of FEM-based structural methods

of analysis for improved accuracy of conceptual-level airframe structure mass estimates - Current effort develops enabling capabilities including

loads, stress analysis/convergence and mass calculation

UT Structural Analysis Module Research Objectives


Some Issues

FEM programs typically require specialized training and experience often not available at CD project level - The issue is not about knowledge but tool specific skills

- Designers understand fundamentals but running specific programs and interpreting specific outputs can get complex - Especially problem definition, file setup and data analysis

Structural loads is another specialty area challenge - Conceptual load cases are often simplistic and don’t capture

key physical environments that end up sizing real structure - Internal load paths can end up being far off the mark

FEM Based Mass Property Estimates Considered Proprietary - Decades of effort but few if any open source publications


Where we started – Manual set-up up of CalculiX files (pages in our first users manual)

a. Mesh import from VSP into CalculiX (3 pages) b. Pre-processing with CGX including trim (7 pages) c. Preparation of loads (5 pages) d. CalculiX solution (1/2 page) e. Model and analysis refinements (6 pages) f. Results import from CalculiX to VSP (2 pages)

Days to weeks of study required just to get started


VSP CalculiX Process – where we went next

Model Center ©

Parametric Thickness

Boundary Conditions

Parametric Loads

Input File Structure to CalculiX

Vehicle Sketch Pad

External and Internal Mesh Generation

Parametric External Geometry

Parametric Internal Geometry

CalculiX ©

FEM Solution

FEM Input

FEM Post Process and Graphics

FEM Output

Model Center ©

VSP Geometry Input

FEM and Geometry Output

Geometry Iteration and Convergence

CurrentPlanned

Model Center ©

Parametric Thickness

Boundary Conditions

Parametric Loads

Input File Structure to CalculiX

Vehicle Sketch Pad




Vehicle Sketch Pad




CalculiX ©

FEM Solution

FEM Input


FEM Output

Model Center ©

VSP Geometry Input

FEM and Geometry Output

Geometry Iteration and Convergence

CurrentPlanned


VSP CalculiX Process – where we are now

UT Input Executable (Java)

Boundary Conditions and Load Cases

CalculiX Input File

Vehicle Sketch Pad




CalculiX ©

FEM Solution

FEM Input


Output Files

UT Convergence Executable (Java)

Solution Files

Thickness Iteration

Stress Convergence

Thickness and Material Properties

Mass Calculation

Wing Trim


Overview of Process

•  Generate wing model in VSP •  Add wing structure in VSP – ribs, spars, skins •  Compute and export mesh •  Run VSP to CalculiX software •  Input required variables in GUI

-  Trim, initial thickness, material properties, load case •  Software runs through CalculiX preparation methods

-  Apply inputs, import mesh, trim wing, define materials, apply load case, fix rib

•  Generate initial CalculiX result -  View stress/strain distributions, deflections, and initial 3D geometry

•  Iterative thickness method -  Calculate new thicknesses based on previous iteration stresses,

thicknesses, and allowable stress -  Run and save each iteration CalculiX analysis

•  Converge on thickness solution •  Compute mass approximations based on converged thicknesses


Background of ACT Wing

•  The Advanced Composite Transport (ACT) aircraft wing is modeled -  We will run a simplified version to demonstrate the software

•  The wing structure is shown below. The wing will be modeled as a single section wing and the LE and TE will be trimmed to analyze the wing structural box only


Wing Structure Modeling - VSP

•  Generate the wing model in VSP using the MS Wing geometry options •  For the ACT Wing, import an ACT wing background image, as shown

-  Note: For the VSP to CalculiX software, the wing must be single section

x

y



•  Define the skins, ribs, and spars using the wing structures feature

Set mesh parameters – Adjust element size and restrictions Adjust for curvature based mesh




Add and delete spars and ribs Select number of

rib/spar to adjust




Adjust properties of ribs/spars/skins – Thickness and density are re-defined in VSP to CalculiX

Compute, export, and show mesh – Set path and name Viewing window showing processes



•  ACT Wing model shown with simplified wing structure (for reduction in run time and processor requirements)

•  Mesh generated for the simplified ACT Wing model


CalculiX FEM Software

Overview • CalculiX is an open-source 3-D finite element method (FEM) program from the Free Software Foundation • CalculiX is an excellent tool, but it is not an overall simple program to use for non-FEM specialist, CD-level users

-  To avoid this issue, UT researchers were able to develop CD-level user-friendly interface using Java scripts that push time-consuming CalculiX specific processes into the background and translate otherwise arcane FEM input requirements into CD user-friendly terms

• CalculiX FEM analyses are used to generate von Mises stress maps as well as displacement, strain, and force plots for CD-level internal wing structural arrangements. • The most significant aspect is that the plots generate by CalculiX are based on high-fidelity engineering methods, and the time expended to generate them is measured in minutes.


CalculiX FEM Software

Method • The mesh generated by VSP is imported to the software and modified in the trim method and boundary conditions are applied restraining the wing motion at a fixed rib from translation in the x-,y-, and z- directions • The external distributed loads are applied at the nodes along a LE, TE, or load spar on the upper or lower surface, and the point loads are applied at the node on the upper or lower surface nearest to the defined location • The material properties are defined for element sets and initial thicknesses at the nodes are defined • CalculiX requires the definitions of mesh geometry, fixed nodes, element material properties, initial nodal thickness, and nodal (point loads) and elemental (distributed loads) load applications

-  Currently only includes the option for elastic material definition, defined by Young’s modulus and Poisson’s ratio

-  The results of the analysis include stresses (principle, von Mises, Tresca), strains, deflections, and external forces

• The iterative thickness re-evaluates the nodal thicknesses based on the previous iteration thicknesses and the CalculiX stresses • CalculiX is re-run with the new thicknesses until the solution has converged


Method and Examples of Trim

Overview •  The trim method is used to trim the leading edge and trailing edge devices

to remove components that are unnecessary for structural analysis -  This simplifies the analysis, reducing the analysis time and simplifies

the wing into just the wing box Method •  Options include trimming the entire leading edge (LE) or trailing edge (TE),

trimming one device on the LE or TE, or trimming two devices on the LE or TE, or no trim for the LE or TE

-  For trimming devices, it is necessary to specify rib numbers between which the trim should be performed

-  The leading edge and trailing edge spar numbers must be specified -  Ribs and spars must therefore exist at the span-wise and chord-wise

locations, respectively, that trim is performed •  Planes are defined along the LE and TE spars and the ribs constraining

the control surfaces to define the wing box and identify the nodes to be deleted



Example of entire leading edge (LE) and trailing edge (TE) trim for ACT Wing - Yields ACT structural wing box



Examples of LE and TE trim with varying numbers of devices:


Method and Examples of Load Cases

Overview • Currently, VSP to CalculiX allows the user to define a load case along a user-defined load spar (must be defined in VSP), point loads on the wing, and along the LE and TE trimmed devices

• Along the load spar, which is typically defined at the quarter chord, a linearly distributed load, elliptically distributed load, or a distributed load defined by Schrenk’s approximation can be defined

-  These methods were chosen because they are simple, commonly used aerodynamic approximations

-  Distributed loads are defined in the vertical direction (z-axis)

• Point loads (forces and moments) are defined by magnitude, direction (Fxx, Fyy, Fzz, Mxx, Myy, Mzz), and location (percent semi-span & percent chord)

• LE and TE trimmed device locations can be loaded with a constant magnitude distributed load or point loads defined by percent span of the device



Linear Load Case Along Load Spar:

The linear load case is defined by the distributed root and tip loads (force per length) input by the user. These loads are then used to calculate approximate point loads for each node along the specified spar. This is accomplished by calculating the equivalent force due to the distributed load from midpoint between the node inboard to the midpoint between the node outboard from the node for which the load is being calculated.



Elliptical Load Case Along Load Spar:

Elliptical loading was used due to the common use and standard practice methods to model a wing in steady level flight. This load is applied at the user-defined loading spar (typically the quarter chord) of the wing, similar to many approximations in accordance with accepted aerospace conventions. In the structural analysis module, the elliptical load case is calculated from aircraft weight and load factor inputs.



Schrenk’s Approximation Along Load Spar:


Iterative Thickness and Mass Generation Methods

Overview: •  The iterative thickness method is used to produce an idealized wing

structure, resulting in a minimum thickness (and therefore minimum weight) solution based on the stress allowable

•  The thicknesses are expected to decrease at the rib and spar webs and increase near the upper and lower surfaces to form an I-beam like section

•  This results in a more rigorous determination of the wing structure mass from physics-based geometry refinement methods

•  Note: The thickness extends inwards and outwards from the mesh surface. This should be taken into account when generating the wing model in VSP

•  Once the solution has converged, the masses for each component type (spars, ribs, and skins) and the total mass are generated from the mesh area, final thickness values, and material density

•  The final masses are displayed in a separate window


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Mass Generation Method


The VSP to CalculiX software GUI is simple and user-friendly. There are three tabs corresponding to trim options and initial thickness, material properties, and load case, shown below. Once all the inputs are entered, hit the “Run” button to start the analysis.

Note: The inputs from the previous run are saved and automatically uploaded into the fields when the program is started. Defaults are used when there is no previous run.

VSP to CalculiX User Interface



Set the directory of VSP mesh files (input and output files)



Set the directory of the CalculiX folder (use default with typical installation)



Specify the file name of the VSP model



Definition of trim – LE and TE devices



Choose to trim LE and TE (yes or no) and input spar number corresponding to the LE and TE – as defined in VSP



Number of devices – 0: Entire LE or TE 1: One device 2: Two devices



Example showing method for trimming two trailing edge devices



If the number of devices is 1 or 2: Select rib numbers constraining trim as defined in VSP



Definition of initial thickness of components



Example of tapered initial thickness input – Tapered span-wise, linearly from root to tip



Define up to four material properties – Requires name, Young’s modulus, Poisson’s ratio, Yield stress, And ultimate stress

Define material for each component – By name, allowable stress, and density, and set minimum gauge and convergence tolerance



Indicate which rib is fixed from translation in the x, y, and z direction – select rib as defined in VSP



Apply the load along the loading spar on either the upper surface or lower surface of the wing



Load case applied along the load spar (typically at the quarter chord)



Choose the load to apply along the loading spar – Linearly distributed, elliptically distributed, Schrenk’s approximation, or none



Select the load spar number as defined in VSP (typically corresponds to the quarter chord)



Input the weight of the aircraft and the load factor for elliptical and Schrenk’s approximation



Input the root and tip distributed loads for the linear load case



Apply external moment about the span-wise axis (y-axis) at the root rib



Define any number of point loads on the upper and lower wing surfaces, clicking “Add” for each point load – defined by magnitude, degree of freedom, and location (percent span and percent chord)



For the point loads case on the LE and TE, define any number of point loads – define by load, degree of freedom, and location (percent device length) and click the “Add” button

For the linear load case on the LE and TE, the constant distributed load along the device is specified only


Example Run-through of ACT Wing

•  VSP Structures •  Inputs for VSP to CalculiX •  Viewing in CalculiX

-  Show stress and strain distributions, displacements -  Show results for each iteration

•  Mass generation results GUI



Verification of CalculiX Results using a VSP wing with a rectangular airfoil (approximating a beam):

The wing is a analyzed as a cantilever rectangular beam loaded along the 50% chord line and fixed at the root rib. The cross sectional geometry is shown below.



Verification of CalculiX results:

Top

Side (Spar)

Root Rib



Verification of CalculiX results:



Verification of mass generation results:


Span Times – ACT Wing

Time in VSP Generate VSP external wing geometry 5 min. Generate VSP internal wing structure 5 min. Generate and export VSP mesh 2 min. Time 12 min. Time in VSP to CalculiX Define boundary conditions 1 sec. Define trim conditions 5 sec. Define material properties 30 sec. Define spar, rib and skin thicknesses 5 sec. Define loads 5 sec. Iterate Calculix solutions 10 min Viewing CalculiX solutions: 15 sec. Mass generation 30 sec. Time 11 min. 31 sec. Total time 23 min. 31sec.


2012 Accomplishments

1.  Implementation of a simplified wing skin trim feature. 2.  Fully automatic input and output file read/write among programs 3.  Standardized wing load cases (linear, elliptical and Schrenk) 4.  Multiple wing load introduction options to include force and

moment introduction along multiple constant percent chord lines 5.  Standardized wing design load cases representative of simple

symmetrical and asymmetric maneuvers 6.  Discrete force and moment point loads to represent landing gear,

engine mounts and nacelles and external pods 7.  GUI based material property and load inputs 8.  Alternate rib pair fixed boundary condition option (i.e. not root rib) 9.  Calculation of structural element thickness required to meet user

defined working level stress, strain or displacement requirements 10. ModelCenter no longer required operating environment

Item 9 was enabling capability for FEM based mass property estimates


Currently Planned Work – Academic Year 2012-13

Design and Analysis (budget and schedule available dependent ) 5.  Distributed fuel and inertia loads including fuel tank pressure

and/or fuel and structural mass inertia loads 6.  More standardized design load conditions including gust loads,

and hard landings 7.  Redefined “point” loads 8.  Buckling defined structure (stringers and other typical features) 9.  Control surface deflection based loads 10.  Parametric conceptual-level pressure (vs. constant chord) loads 11.  Effects of variable structural "contact" definition

Applications and Comparisons (focus for the year) 1.  Advanced Composite Technology (ACT) Wing Comparison

(stress and mass) 2.  X-56A Wing (stress and mass comparison) 3.  Empennage structure (horizontal and vertical, mass estimation) 4.  NASA TM 110392 wing weight comparisons (from parametrics) 5.  NASA Langley Workshop


FY 12 Deliverables

•  Developed Codes •  Structural Module Users Guide •  User Workshop Presentation


Future Work

With VSP Development Team •  Multi-section wings •  Specialized user defined loads •  Design superposition of multiple load cases •  Video-based documentation •  Links to other modules (e.g. aero loads) •  Multiple applications and mass calibrations •  NASA Langley User Workshop

With UT Arlington •  Effects of FEM model simplification (especially buckling) •  Realistic CD/PD-level solutions for structural design features left

out of FEM analysis model •  Composite structure •  Fuselage and nacelle structure

With Other Government and Industry Collaborators •  Proprietary airframe comparisons •  Non-primary airframe load carrying structure and effects


Questions


Notes on Functionality

•  The iterative results tend to be more stable when starting with an excessively thick wing (using higher than expected values for initial thickness declarations) such that the thickness tends to reduce.

•  The iterative solutions tend to be more stable for low loading, low minimum gauge and high loading, high minimum gauge.

•  Check the mesh generated by VSP before continuing with a run. Sometimes a bad mesh is generated, as shown below, which will cause an error in the run. The issue is typically solved by decreasing the element size and re-meshing the wing.

•  Specifying a fine tolerance will increase run-time and can potentially result in the solution never converging.

•  Depending on the computer, a fine mesh on a complex wing may cause issues for running the software due to memory allocation failure. If there is a problem, check the file size of the mesh files exported from VSP. Problems tend to start for mesh geometry file sizes greater that 5,000 KB.

•  High complexity in the model also tends to significantly increase the runtime.

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