VSP to CalculiX User Interface

Transcription

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
© 2012 Armand J. Chaput
UT Structural Analysis Module Research Objectives
(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
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
Vehicle Sketch Pad
Model Center ©
Parametric
External Geometry
Boundary Conditions
Parametric
Internal Geometry
Parametric Thickness
Parametric Loads
External and Internal
Mesh Generation
Input File Structure
to CalculiX
Model Center ©
CalculiX ©
Geometry Iteration
and Convergence
FEM Input
VSP Geometry Input
FEM and Geometry
Output
Planned
FEM Solution
FEM Post Process
and Graphics
FEM Output
Current
VSP CalculiX Process – where we are now
Vehicle Sketch Pad
Parametric
Internal Geometry
Parametric
External Geometry
External and Internal Mesh
Generation
UT Convergence Executable
(Java)
Thickness Iteration
Solution Files
UT Input Executable (Java)
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
CalculiX ©
FEM Input
FEM Solution
Stress Convergence
FEM Post Process
and Graphics
Mass Calculation
Output Files
Overview of Process
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• 
• 
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• 
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
y
x
•  Define the skins, ribs, and spars using the wing structures feature
Set mesh parameters –
Adjust element size
and restrictions
Adjust for curvature
based mesh
Select number of
rib/spar to adjust
Add and delete
spars and ribs
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 userdefined 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.
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 material
for each
component –
By name,
allowable stress,
and density, and
set minimum
gauge and
convergence
tolerance
Define up to four
material
properties –
Requires name,
Young s
modulus,
Poisson s ratio,
Yield stress,
And ultimate
stress
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 linear
load case on
the LE and TE,
the constant
distributed load
along the
device is
specified only
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
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:
Side (Spar)
Top
Root Rib
Verification of mass generation results:
Span Times – ACT Wing
Time in VSP
Generate VSP external wing geometry
Generate VSP internal wing structure
Generate and export VSP mesh
Time
5 min.
5 min.
2 min.
12 min.
Time in VSP to CalculiX
Define boundary conditions
Define trim conditions
Define material properties
Define spar, rib and skin thicknesses
Define loads
Iterate Calculix solutions
Viewing CalculiX solutions:
Mass generation
Time
1 sec.
5 sec.
30 sec.
5 sec.
5 sec.
10 min
15 sec.
30 sec.
11 min. 31 sec.
Total time
23 min. 31sec.
2012 Accomplishments
1. 
2. 
3. 
4. 
Implementation of a simplified wing skin trim feature.
Fully automatic input and output file read/write among programs
Standardized wing load cases (linear, elliptical and Schrenk)
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
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
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
FY 12 Deliverables
• 
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• 
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.

VSP to CalculiX User Interface

Transcription

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