university of texas vsp structural analysis module update...

© 2013 Armand J. Chaput

University of Texas VSP Structural

Analysis Module Update - Overview

2nd VSP Workshop, San Luis Obispo, CA

Armand J. Chaput, Principal Investigator

Hersh Amin, Undergraduate Research Assistant

Department of Aerospace Engineering and

Engineering Mechanics, University of Texas at Austin

7 August 2013

Dr. Armand Chaput, Department of Aerospace Engineering and

Engineering Mechanics (ASE/EM), University of Texas at Austin (UT)

• Director, Air System Laboratory (2008 - present)

• Senior Lecturer, Air System Engineering Design (UAS focus)

• Lockheed Martin – 30 years, advanced product development

• Air System Design and Integration - Senior Technical Fellow

including assignment as F-35 Joint Strike Fighter Weight “Czar”

• Unmanned Combat Air Vehicles - Integrated product team lead

• National Aerospace Plane (NASP) - National team Chief Engineer

• Advanced Design Department – Manager

Undergraduate Research Assistant Team

Hersh Amin, Research Ass’t

Josh Eboh, Team Lead

Natalie Maka , Research Ass’t

Patil Tabanian, Research Intern

Self Introductions

Based on work performed under NASA/NIA Task Order 6322-UTEX,

“Advanced Conceptual Design Tools and Development”

Armand J. Chaput 2013

Sarah Brown, Research Ass’t

Jose Galvan, Research Ass’t

Alex Haecker, Research Ass’t

Tejas Kulkarni, Team Lead

Current Previous


(1) Expand VSP user capabilities for employing higher

order, physics based tools and methods during

conceptual design (CD)

(2) Integrate VSP FEA structures module with an open

source finite element method (FEM) structural analysis

program in a user friendly interactive environment

- Currently focused on CalculiX (available under terms of GNU

General Public License as published by the Free Software Foundation)

(3) Develop basic capabilities for open source application

of FEM-based mass property (MP) methods to CD

- Current effort develops and validates fundamental wing

and tail mass estimation methodologies

- Follow-on effort (?) will expand applications

VSP Structural Analysis Module (SAM) R&D Objectives


2012 VSP Workshop: Version 0 (VSP structures module

integrated with CalculiX, posted Sept 2012)

- UT Java scripts simplify setup and run unitary VSP FEM model

- GUI inputs (loads, constraints, material prop, trim to wing box)

- CalculiX solution and display of stress, strain, displacement

- Calculate mass of input FEM model

- Initial stress/mass results (convergence stability issues)

May 2013: Version 1 (fully stressed coarse grid FEM & mass)

- Separate upper and lower skin, spar, rib FEMs connected by

rigid body nodes resolved convergence issues

- Adds: skin section trim (1.0) and solution status feedback (1.1)

Today: Version 2 (inertia loads and sizing for multiple load cases)

- Angle of attack plus fuel and discrete mass inertias

- Convergence for multiple load cases

- Initial calibration/validation results

Objective 1 - Expanded Capabilities

Codes and users guide posted

at: http://vspsam.ae.utexas.edu/


Objective 2 - CalculiX Integration

UT Input Executable (Java)

Boundary Conditions and Load Cases

CalculiX Input File

Vehicle Sketch Pad

External and Internal Mesh

Generation

Parametric

External Geometry

Parametric

Internal Geometry

CalculiX

FEM Solution

FEM Input

FEM Post Process

and Graphics

Output Files

UT Convergence Executable

(Java)

Solution Files

Thickness Iteration

Stress Convergence

Thickness and Material Properties

Mass Calculation

Wing Trim


Last iteration –

thickness “converged”

yellow = 28.9 ksi

2 spars,

5 ribs

Last iteration –


yellow = 28.7 ksi

2 spars,

7 ribs

Last iteration –


yellow = 28.4 ksi

5 spars,

15 ribs

Objective 3 – FEM mass for fully stressed trade study

wings with minimum gage constraints (from last year)

• Lightly loaded notional wing

• VSP defined spars and ribs

• VSP SAM defined thickness, materials

and loads (2D running load along 0.25c)

• 30 ksi fully-stressed design objective

• FEM mass calculated

• Time to generate-solve-converge for all 3

solutions from scratch < 3 hr

• Issue – solution stability


So What’s New?

• FEM structural methods have been available for decades

- FEM analysis requires a well-defined representation of the

airframe structure; design details are not available during CD

- CD design and analysis cycles are typically incompatible with

time required for FEM model development and turn around

- By the time a FEM model is developed the CD team has

usually moved on to another concept

- CD budgets are often incompatible with specialized FEM

analyst staffing requirements

• VSP SAM lets structural design and analysis keep up with

other CD participants

- Traditional FEM model definition, solution and analysis time

and skill requirements limit wide scale application

VSP SAM enables requirement-based CD mass estimation


Background – Airframe Mass Property (MP) Estimation

Airframe mass is driven by multiple requirements; many of

which are not captured by traditional CD analysis methods

- Current state-of-the-art MP methods still rely on parametric

(statistical or regression analysis of historical data) methods

- A problem when trying to predict mass for new vehicles, new

materials, new processes or new design requirements

Primary loads drive 60% of load carrying airframe mass

- Calibrated FEM analyses should be able to predict primary

structural mass with better accuracy than parametrics

Secondary structure mass is driven by non-primary loads

- Many of which could be captured by FEM-based methods

System installation and integration effects are problematic

- Not defined until much later in the design process

Bottom line: FEM-based methods can improve the quality

of at least CD and PD primary structure mass estimates

- It doesn’t cost any more or take any more time


FS 228.2

FS 283.9

0.70 c?

182.6

0.83 c

0.15 c

318”

117

”

MP Data from Grumman Aerospace

A-6E Weight Report

WT-128R-1S37 Aug 1988

Courtesy of Paul Kachurak, NAVAIR

Why we need improved CD methods - Example from UT method development effort (A-6E)

1. Raymer (fighter-attack):

= 0.0103 [(Wdgnzdu)0.5 Sref

0.622 AR0.785

(1+)0.05 Scsw0.04]] / [(t/c)0.4

Cos(0.25c)]

= 4092 lbm

2. Nicolai (USN fighter):

= 19.29 [(Wtonzdu)/(t/c)] {[(Tan le -

2(1-) / AR(1+)] 2 +1] 10-6} 0.464

[(AR(1+ )] 0.70Sref0.58 = 7057 lbm

A-6E WING GROUP (lbm) Sum

WING STRUCTURE - BASIC 3443 3443

SECONDARY STRUCTURE 931 4374

TRAILING EDGE DEVICES 593 4966

LEADING EDGE DEVICES 241 5207

SPEED BRAKES 145 5352

WING GROUP - TOTAL 5352

Inc. wing fold unique 297

Typical CD wing parametric estimates:

Wdg = 36526 lbm

Wto (land) = 60705 lbm


Perspective – Why airframes weigh what they weigh

For a good design, the driver is structural requirements

- Operating environment (speed, altitude and temperature)

- Almost always known and available up front

- Failure modes, Durability and Damage Tolerance (DaDT)

- Loads inc. primary air loads, secondary loads and accidents

- All are quantifiable but often missed in CD (inexperience)

- Systems integration (loads, penetration, installation access)

- Predictable but only when design teams are integrated

- In-flight moving parts (control surfaces, doors, gaps & locks)

- Ground handling and maintenance access

- Manufacturing and assembly (including workforce skill level)

- LCC cost, schedule, risk and growth considerations

During early phases, many designers use rule of thumb or

program defined knock-downs to cover unknowns

- Generally expressed in terms of % design stress (or strain)

- Similar to our “Conceptual Design Nominal Stress (CDN )”

much k

now

n

much u

nkno

wn


FEM Approach to Nominal Stress (CDN) for CD

Step 1 - Develop CD-Level FEM Models of Existing Designs

- Capture representative geometry, material and primary loads

- Focus on primary structure: Spars, Ribs and Skins,

Step 2 - Back out CDN to correlate calculated FEM mass with

actual mass consistent with min gage

1. Repeat Process for Multiple Vehicles in Given Class

2. Use Correlated CDN for CD and Early PD Designs

B747 Simplified Rib Model before & after Trim

Iter

Spar Mass (lbm)

Skin Mass (lbm)

Rib Mass (lbm)

Total Mass (lbm)

1 13392 26457 6249 46100

2 10506 21179 3750 35435

3 8984 18572 2472 30030

4 8157 17315 1817 27290

5 7677 17189 1481 26348

6 7280 17022 1308 25610

7 7023 17238 1221 25482

Solution for DNS = 46.5 Ksi


1. Solve simple problems first (current effort)

• Geometry - trapezoidal wing box, equivalent skin thickness

• Loads - symmetrical pull-up, push-over, 2-D distributed loads

(inc. Schrenk approx.) with user defined spar load fraction

• In-plane isotropic properties – assumed for simplicity

• No fasteners or other non-optimums - i.e. “knocked-down”

static stress sized structural representation of a real wing

2. Address buckling as separate issue using CalculiX

buckling factor

3. Next (?) - apply design “rules of thumb” to estimate

basic non-optimums (fasteners and spacing, fuel tanks and

sealant, load introduction fittings, hinges, tracks, etc.)

CDN Mass Estimation Methodology Strategy

Proposed methods are open source and publically available

to encourage collaborative development approach


Good/Bad News - CD Nominal Stress (CDN ) Method

Pro

• Simple and straight forward, anybody can do it

• Based on physics, geometry and CD design requirements

• Based on familiar structural tools and methods

• Reduces reliance on Fudge Factors for MP estimates

• Can be tailored for internal capabilities and skills

• Good risk tracking metric for customer engineers

Con

1. Little CD history

2. Limited publically available correlation data

3. Methodology for Non-Primary Loads, Fittings and Fasteners

and System Integration needs development


Transport Wings

• Advanced Composite Technology (ACT) Test Wing

- PDF sketch level geometry, acceptable MP detail

- Materials defined, no allowables (nominal GrEP assumed)

• B707, B727, B737 (early models)

- Based on 1990s NASA supported PDCyl published data

- PDF sketch level geometry, no detail below wing box level

- Materials assumed (nominal 2024 T3 )

A-6E Wings (metal and composite replacement)

• Biz-jet type planform (exc. for folding wing)

- Good MP data quality, good for Version 2 validation

- PDF sketch level geometry plus handbook data

- Materials assumed (nominal 2024 or GrEP)

UAV Wings (in progress)

• X-56A Wing (1 of 4)

- PDF based geometry, good MP, nominal GrEP assumed

UT Methodology Development/Calibration Status


Advanced Composite Technology (ACT) Static Test

Article (as reconstructed and analyzed) Mass Categories lbm Fraction

Upper and lower skins 1656 0.41

Spar caps and stringers 1238 0.31

Spar webs 350 0.09

Stress based (subtotal) 3244 0.81

Aero ribs and intercostals 520 0.13

MLG rib blkhd 75

MLG pad up 31 0.03

Bolts and nuts 80 0.02

SOB Pad up 25

Access panel pad-up 45 0.02

Total 4020 1.00

Mass analysis based on data from NASA

CR-2001-210650-AST Composite Wing

Program-Executive Summary

Reconstructed mid-

chord thickness

Load actuatorsLoad actuators

LE 30

Stringer runout

Stringer runout

Fiber – IM7 and AS4

Process - GrEP VARTM

Sources: multiple NASA and Boeing ACT Documents

ACT Transport

ACT Test Article

ACT thickness

not linear

Actuator

load (lbf) 2y/b

1 40500 0.947

2 99750 1.000

3 -3000 0.634

4 21000 0.704

5 15000 0.384

6 -45000 0.382

7 45000 0.202

8 6000 0.296

Wing carry-

through not

included in

ACT test box

Spar web CDN = 27.0 ksi

Ribs CDN = 37.9 ksi

Skin CDN = 36.5 ksi

Min gage (given)

= 0.22 in


Generic Boeing Transport Wings – B727, B737, B747

Aircraft or

article WDG [lbs]

DUL

Nz

Half

span Sref

[ft2]

Sweep

[deg]

Half

span [ft]

Half

span

AR TR (t/c)r (t/c)t

Dihed

[deg]

Wing

Fuel

Engine

pylon or

store (per

side)

Engine or

store

mass (ea)

747-100 713,000 3.75 2790 37.5 98.5 6.96 0.265 0.1794 0.078 7 Y 2 8608

737-200 100,800 3.75 502.5 25 45.4 8.21 0.220 0.126 0.112 6 Y 1

727-300 160,000 3.75 793.5 32 55.2 7.67 0.265 0.154 0.09 3 Y 0

ACT Test unk 3.75 N

A-6E 43,077 9.75 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 N 0 N

A-6E 43,077 9.75 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 Y 2 2296

A-6E 43,077 -4.5 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 N 0 N

NASA 110392

Wing Weights

PDCYL

(lbm)

Load-

carrying

structure

(lbm)

Primary

structure

(lbm)

Total

structure

(lbm)

B-727 8688 8791 12388 17860

B-737 5717 5414 7671 10687

B-747 52950 50395 68761 88202

B-720 13962 11747 18914 23528

DC-8 22080 19130 27924 35330

MD-11 33617 35157 47614 62985

MD-83 6953 8720 11553 15839

L-1011 25034 28355 36101 46233

B747-100 VSP

Simplified to 15 ribs

B747, Min gage =

0.1” exc. ribs = 0.36”

CDN = 46.5 ksi

B727-100 VSP

Full 26 ribs B727, Min

gage = 0.1”

CDN = 27.4 ksi

B737-100 VSP

Simplified to 14 ribs B737, Min gage =

0.1” exc ribs = 0.15”

CDN = 29.8 ksi

Parametric geometry from Analytical Fuselage and Wing Weight Estimation of Transport Aircraft, NASA TM 110392, May 1996

Wing Boxes Notional and Extrapolated to Centerline


Reconstructed A-6E Geometry

All locations and linear

dimensions in inches

Wing Box (theo)

b/2 = 305”

Cr (theo) = 117”

Ct = 33”

Theo taper ratio = 0.282

Taper (fold) = 0.661

Sref = 317.7 sqft

Aspect ratio = 8.13

t/c (BL0) = 0.143

t/c (BL144) = 0.137

t/c (BL 305) = 0.115

Kc = 1.0

Theoretical Wing

b/2 = 318”

Cr (theo) = 182.6”

Ct (theo) = 57”

Taper ratio = 0.312

Sref = 528.9 sqft

Aspect ratio = 5.31

Taper ratio = 0.312


t/c (BL33) = .09

t/c (BL144) = 0.084

t/c (tip) = 0.059

LE flap = 15% c

Flaperon LE at 65% c

Est. box Kc = 0.5

BL

30

5

BL

31

8

FS 228.2

FS 283.9

33”

BL

0

0.70 c?

182.6

BL

14

4

FS 0

56”

0.83 c

0.15 c

57”

318”

B

L 7

8

B

L 6

6

0.05 c

0.70 c

28.5

28

117

”

• Wing fuel – L/R full 6923 lbm

• 2 Ext fuel inbd – L/R 4010 lbm (2)

• 2 Tank w/adapter inbd L/R 398 lbm

• 2 Pylon inbd L/R : 192.6 lbm (WS 95)

• 2 Ext fuel outbd – L/R 4010 lbm

• 2 Tank w/adapter outbd L/R 398 lbm

• 2 Pylon outbd L/R : 183.4 lbm (WS 141)

• Pylon (WS 187) – replacement wing only

• Max GW landplane: 60705 lbm

• Max t/o landplane: 60400 lbm

• Flt des GW (landplane): 36526 DUL = 9.75

• Max GW zero fuel, zero stores = 39781 (body

fuel =9016 lbm)

BL

38

.9

Wing Box (less ctr sect)

b/2 = 239”

Cr = 98.8”

Ct = 33”

Theo taper ratio = 0.334


Sref = 218.8 sqft

Aspect ratio = 7.25

t/c (BL66) = 0.140

t/c (BL144) = 0.137

t/c (BL 305) = 0.115

Kc = .5/.65 = 0.77

0.15 c

0.65 c

A-6 graphic and raw data from A-6

Post Design Analysis Report (CORL

G003), Dec 1993


A-6E MP Summary Data and VSP SAM Results

JS&F = Joints, splices and fasteners

TipFaFen = tips, fairings and fences

Wing box

mat'l Skins

Skin

stiffners

Spars and

stiffeners Ribs Rib Bkhd Js&F TipFaFen Other Pri - sum TipFaFen Doors Acs Panel Fin & walk Misc Sec - sum CV Unique Total

Wing primary - lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm

CENTER SECTION 718.00 0.00 39.50 0.00 103.60 35.40 0.00 3.40 899.90 1799.80

INTERMEDIATE PANEL 1077.80 681.40 0.00 200.00 0.00 196.40 106.10 0.00 0.00 1183.90 2367.80

OUTER PANEL 814.70 488.30 1.30 239.50 44.20 41.40 58.90 0.00 0.00 873.60 1747.20

LE 142.70 86.30 13.30 15.40 27.70 6.40 0.00 0.00 149.10 298.20

TE 68.60 37.40 0.00 19.20 12.00 0.00 15.20 0.00 83.80 167.60

Wing and intergration 35.8 35.80

WING TOTAL 2964.90 2011.40 14.60 513.60 83.90 341.40 206.80 51.00 3.40 3190.30 231.80 165.90 2.00 125.60 3.80 529.10 295.70 4015.10

Less center section 2103.80 1293.40 14.60 474.10 83.90 237.80 171.40

SECONDARY STRUCTUREPRIMARY STRUCTURE

Data from: MODEL A-6E ACTUAL DETAIL WEIGHT AND

BALANCE REPORT, NO. WT-128R-1S37, August 1988


Ribs CDN = 12.4 ksi

Skin CDN = 46.2 ksi

A-6E @ Wfdg = 36526 lbm

no ext tanks, no wing fuel,

Min gage = 0.040”

A-6E @ Wfdg = 36526 lbm with

2x2300 lbm ext tanks + 3462 lbm

wing fuel, Min gage = 0.040”


Ribs CDN = 9.0 ksi

Skin CDN = 42.6 ksi

Note - Rib definition includes bulkheads and store stations

- Spar web fuel pressure, cat. and arrest loads not included

Overall results correlate with expectations

- Additional load cases and refined geometry expected to improve results


VSP SAM CDN Preliminary Results

CD Nominal Sress Correlation

0

5

10

15

20

25

30

35

40

45

50

0 200 400 600 800

WL x Nz (psf)

Sig

ma

CD

N (

ks

i)

A-6 spweb

A-6 rib

A-6 skin

ACT spweb

ACT rib

ACT skin

B727-300

B737-200

B747-100

A-6E weighted

ACT weighted

Overall wing box mass weighted results


VSP SAM CDN Issue – VSP Geometry Constraint

A-6E Wing Thickness Profile

0.05

0.08

0.10

0.13

0.15

0 50 100 150 200 250 300 350

BL (in)

t/c

WingTheo boxTheo box less ctr

A-6E Wing Chord Distribution

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 50 100 150 200 250 300 350

BL (in)

t/c

Wing

Wing box

Actual t/c distribution

Assumed t/c distribution

0.00

5.00

10.00

15.00

20.00

25.00

30.00

35.00

40.00

45.00

0 100 200 300 400 500 600

ht

(in

)

ACT Wing Box t/c

Least Squares Fit

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0 100 200 300 400 500BL (in)

t/c

Max to min t/c

VSP structural module grid limited to single trapezoidal planform and stream-wise chord

- Even CD level structures need more flexible multi-panel capability - Effect on ACT

- In order of priority we need (1) non-linear t/c, (2) non-linear chord and (3) non-stream-wise cut capabilities

Case 1 Case 2 Diff

Skin 36.46 47.31 30%

Ribs 37.85 31.94 -16%

Spars 27.02 34.04 26%

CDNS (psi)


Concluding remarks - VSP Structural Design and Analysis

1. FEM design and analysis can be accommodated during

conceptual design without adding onerous requirements

for higher levels of design detail

- Serious structural design issues can be assessed, identified

and resolved without slowing down the pace of design

2. Design and analysis methods can be applied intelligently

by designers who are not structural specialists

- Tool specific details can be pushed into the background

3. FEM model results correlate with actual mass property

data through use of data validated CDN

4. VSP SAM based methods are now ready to replace

parametric mass estimate methods for primary load

carrying wing structure


Future Plans

1. Contract nearing completion – tasks to complete in

order of priority

a. Document results, present at 2014 AIAA ASM

b. Complete Version 2 validation

c. Post Version 2 Software and Users Guide

d. Additional methodology calibration


Version 2 Overview - New Features



CalculiX Input File

Vehicle Sketch Pad

External and Internal Mesh

Generation

Parametric

External Geometry

Parametric

Internal Geometry

CalculiX

FEM Solution

FEM Input

FEM Post Process

and Graphics

Output Files

UT Convergence Executable

(Java)

Solution Files

Thickness Iteration

Stress Convergence


Mass Calculation

Wing Trim


Wing Trim – Structural Model Geometry



CalculiX Input File


Wing Trim • Deletes non-primary load carrying structure

- Typically leading and trailing edge devices

• Deletes non-load carrying skin panels

- To represent typical fabric or film skin sections


Inertia Loads (relief)



CalculiX Input File


Wing Trim

Front view – Notional Wing

Fuel mass

Engine and pylon

External store

• Discrete loads applied to rib and/or spar centroid at defined rib and/or spar

• Fuel inertia applied as internal pressure load along bottom (+nz) or top (-nz)

of defined rib/spar tank boundaries

nz


Multiple Load Cases



CalculiX Input File


Wing Trim

Adapted from http://upload.wikimedia.org/wikipedia/commons/1/13/PerformanceEnvelope.gif

1 2

4 3

1 – Max + 2 - Min +

4 - Max - 3 - Min -

Multiple load case methodology sizes structure (and

calculates mass) for most demanding of multiple cases


Running CalculiX

Plus user feedback on solution status


CalculiX Buckling Factor (BLF)

CalculiX Linear Buckling Analysis

0.60

0.70

0.80

0.90

1.00

1.10

1.20

4 8 12 16

Number of Ribs

Bu

cklin

g Fa

cto

r (B

LF)

CalculiX BLF output can provide design guidance on

spacing for stability but at cost of 2x solution time

CalculiX

FEM Solution

FEM Input

FEM Post Process

and Graphics

Output Files


Questions


Reconstructed mid-

chord thickness

Actuator # Rib # Location

TE sta.

fm CL bte

C perp to

TE (in)

Chord

fraction

from TE

BL from

SOB (in)

Max span

fraction

Chord

fraction Load (lbf)

1 18 TE 543.5 453.2 59.42 0 410.7 0.942 0 27000

2 18 LE 543.6 453.3 59.42 1 435.9 1.000 1 66500

3 13 TE 396.5 306.2 72.86 0 277.5 0.637 0 -2000

4 13 LE 396.5 306.2 72.86 1 308.3 0.707 1 14000

5 9 TE 279.5 189.2 83.55 0 171.5 0.393 0 10000

6 7.50 LE 236.75 146.45 87.46 1 169.7 0.389 1 -30000

7 6 TE 194 103.7 91.37 0 94.0 0.216 0 30000

8 6 LE 194 103.7 91.37 1 132.6 0.304 1 4000

Actuator # Rib # Location

TE sta.

fm CL bte

C perp to

TE (in)

Chord

fraction

from TE

BL from

SOB (in)

Max span

fraction

Chord

fraction Load (lbf)

1 18 TE 543.5 453.2 59.42 0 410.7 0.942 0 27000

2 18 LE 543.6 453.3 59.42 1 435.9 1.000 1 66500

3 13 TE 396.5 306.2 72.86 0 277.5 0.637 0 -2000

4 13 LE 396.5 306.2 72.86 1 308.3 0.707 1 14000

5 9 TE 279.5 189.2 83.55 0 171.5 0.393 0 10000

6 7.50 LE 236.75 146.45 87.46 1 169.7 0.389 1 -30000

7 6 TE 194 103.7 91.37 0 94.0 0.216 0 30000

8 6 LE 194 103.7 91.37 1 132.6 0.304 1 4000

Load actuators

LE 30

Stringer runout

Stringer runout

Fiber – IM7 and AS4

Process - GrEP VARTM

Sources: multiple NASA and Boeing ACT Documents

ACT Transport

ACT Test Article