C-27J NASTRAN Global Finite Element Model User Manual C-27J … · 2018-05-09 · UNCLASSIFIED. UNCLASSIFIED. C-27J NASTRAN Global Finite Element Model User Manual C-27J-DST-v3.0

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C-27J NASTRAN Global Finite Element Model User Manual C27J-DST-v3.0

M. Opie and D. Hadcroft1

Aerospace Division

Defence Science and Technology Group

1QinetiQ AeroStructures

DST-Group-TN-1700

ABSTRACT The Royal Australian Air Force (RAAF) has commenced a fleet acquisition of C-27J aircraft (AIR8000 Phase 2) to support RAAF tactical airlift capability requirements. As part of the Structural Substantiation Program, a global Finite Element Model (FEM) of the C-27J airframe was obtained from the original equipment manufacturer Alenia Aeronautica. A global airframe FEM is an important supplementary tool in support of current and future RAAF C-27J structural integrity management. In the present user manual, detailed descriptions of the various sub-models constituent to the global model, including guidelines on their use are provided. Pending experimental validation, the enhanced and verified C-27J Global FEM is a linear elastic internal loads model that will be a useful tool in providing global loads results such as wing tip displacements, field stresses, running loads, connection forces, and other potential uses as described within the body of this document. The user manual provides important information on how to use and update the model, and the limitations associated with it.

RELEASE LIMITATION Approved for Public Release

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Produced by Aerospace Division Defence Science and Technology Group 506 Lorimer St Fishermans Bend, Victoria 3207 Australia Telephone: 1300 333 362 Commonwealth of Australia 2017 November 2017 AR-017-020

APPROVED FOR PUBLIC RELEASE

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C-27J NASTRAN Global Finite Element Model

User Manual C-27J-DST-v3.0

Executive Summary

The Royal Australian Air Force (RAAF) has commenced a fleet acquisition of C-27J aircraft (AIR8000 Phase 2) to support RAAF tactical airlift capability requirements. As part of the Structural Substantiation Program, a global Finite Element Model (FEM) of the C-27J airframe was obtained from the Original Equipment Manufacturer (OEM) Alenia Aeronautica. In 2013, DST Group completed a preliminary review of the C-27J Global FEM and recommended several areas of model enhancement, including specific C-27J FEM verification and validation activities. The objectives of these enhancements were: alignment with DST Group model orientation norms, improved useability, alignment with design baseline data, and a reduction of potential usage errors.

A global airframe FEM is an important supplementary tool in support of current and future RAAF C-27J structural integrity management. Previously, no comprehensive FEM user manual or model verification process existed. Hence, in the present user manual, detailed descriptions of the various sub-models constituent to the global model, including guidelines on their use, are provided. Each sub-model is allocated a particular numbering regime, with space for additional nodes and elements, thereby allowing sub-model modification and augmentation that can be easily spliced into the global FEM. Pending experimental validation, the enhanced and verified C-27J Global FEM is a linear elastic internal loads model that will be a useful tool in providing global loads results such as wing tip displacements, field stresses, running loads, and connection forces, as well as providing a foundation for more detailed stress analysis. However, assessment has also identified some remaining limitations associated with legacy modelling constructs which are also discussed within this manual.

This user manual provides critical information on how to use and update the C-27J Global FEM. The limitations associated with this model require focussed consideration by users; it is vital that all users of the C-27J Global FEM are familiar with this manual before working with the model.

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Contents

1. INTRODUCTION ............................................................................................................... 1

2. BACKGROUND .................................................................................................................. 2

3. MODEL ENHANCEMENTS AND QUALITY CHECKS ............................................ 3

4. HOW TO USE THIS FEM ................................................................................................. 4 4.1 Fundamental Principles ........................................................................................... 4 4.2 Limitations ................................................................................................................. 4 4.3 Modifying the FEM .................................................................................................. 4 4.4 Support to Aircraft Structural Integrity ................................................................ 6 4.5 Use as a Loads or Gross Remote Stress Model .................................................... 6 4.6 Use as a Stress Model ............................................................................................... 6 4.7 Support to Detailed Stress Analysis ..................................................................... 6

4.7.1 Local refinement using 3D solid elements ........................................... 6 4.7.2 Derivation of interface loads at sub-model boundaries ..................... 7 4.7.3 Aid aircraft development ....................................................................... 7 4.7.4 Cross-check stress transfer functions by deriving scale factors ........ 7 4.7.5 Damage tolerance analysis ..................................................................... 7 4.7.6 Determine stress transfer functions ...................................................... 7

4.8 Support to Aircraft or Aircraft Component Testing ........................................... 8 4.8.1 Strain, stress, load and displacement predictions ............................... 8 4.8.2 Guidance for strain gauge placement on test articles ........................ 8 4.8.3 Validation of test representativeness .................................................... 8 4.8.4 Virtual ground calibration ..................................................................... 8

5. AIRCRAFT CONFIGURATION AND MODELLING PHILOSOPHY .................... 9 5.1 Wing ............................................................................................................................ 9

5.1.1 Leading and Trailing Edge components .............................................. 9 5.1.2 Ribs ............................................................................................................ 9 5.1.3 Covers ..................................................................................................... 10

5.2 Forward Fuselage .................................................................................................... 10 5.3 Centre Fuselage ....................................................................................................... 11 5.4 Aft Fuselage ............................................................................................................. 12 5.5 Empennage ............................................................................................................... 13 5.6 Landing Gear Attachments ................................................................................... 13 5.7 Aircraft Configuration – Summary ...................................................................... 13

6. LOAD CASES .................................................................................................................... 15

7. FEM NUMBERING SYSTEM ......................................................................................... 18


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8. MATERIAL PROPERTIES AND UNITS OF MEASURE ......................................... 25

9. COORDINATE SYSTEMS .............................................................................................. 26

10. STRUCTURAL IDEALISATION AND SUB-MODEL DESCRIPTIONS ............. 31 10.1 Forward Fuselage .................................................................................................... 31

10.1.1 General arrangement ............................................................................ 31 10.1.2 Shell elements......................................................................................... 32 10.1.3 Bar elements ........................................................................................... 33 10.1.4 Materials ................................................................................................. 34

10.2 Centre Fuselage ....................................................................................................... 35 10.2.1 General arrangement ............................................................................ 35 10.2.2 Shell elements......................................................................................... 36 10.2.3 Beam elements ....................................................................................... 36 10.2.4 Material allocation ................................................................................. 37

10.3 Aft Fuselage ............................................................................................................. 38 10.3.1 General arrangement ............................................................................ 38 10.3.2 Shell elements......................................................................................... 39 10.3.3 Beam elements ....................................................................................... 39 10.3.4 Material allocation ................................................................................. 39

10.4 Centre Wing ............................................................................................................. 40 10.4.1 General arrangement ............................................................................ 40 10.4.2 Shell elements......................................................................................... 41 10.4.3 Beam elements ....................................................................................... 41 10.4.4 Material allocation ................................................................................. 42

10.5 Outer Wing (starboard side shown, port side by similarity) .......................... 42 10.5.1 General arrangement ............................................................................ 42 10.5.2 Shell elements......................................................................................... 43 10.5.3 Beam elements ....................................................................................... 43 10.5.4 Material allocation ................................................................................. 43

10.6 Vertical Stabiliser and Rudder ............................................................................. 44 10.6.1 General arrangement ............................................................................ 44 10.6.2 Shell elements......................................................................................... 44 10.6.3 Beam elements ....................................................................................... 45 10.6.4 Material allocation ................................................................................. 46

10.7 Horizontal Stabiliser (starboard shown; port by similarity) .......................... 47 10.7.1 General arrangement ............................................................................ 47 10.7.2 Shell elements......................................................................................... 47 10.7.3 Beam elements ....................................................................................... 48 10.7.4 Material allocation ................................................................................. 48

10.8 Horizontal Stabiliser – Elevator ........................................................................... 49 10.8.1 General arrangement ............................................................................ 49 10.8.2 Shell elements......................................................................................... 49 10.8.3 Beam elements ....................................................................................... 50 10.8.4 Material allocation ................................................................................. 50


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10.9 Cargo Door ............................................................................................................... 51 10.9.1 General arrangement ............................................................................ 51 10.9.2 Shell elements......................................................................................... 52 10.9.3 Beam elements ....................................................................................... 52 10.9.4 Material allocation ................................................................................. 53

10.10 Cargo Ramp .............................................................................................................. 54 10.10.1 General arrangement ............................................................................ 54 10.10.2 Shell elements......................................................................................... 55 10.10.3 Beam elements ....................................................................................... 55 10.10.4 Material allocation ................................................................................. 56

11. CONCLUSIONS AND RECOMMENDATIONS FOR FUTURE WORK ............. 57

12. REFERENCES .................................................................................................................... 58


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Glossary

ADF Australian Defence Force

ASIP

ASIMP

Aircraft Structural Integrity Program

Aircraft Structural Integrity Management Plan

BDF Bulk Data File

DST Group Defence Science and Technology Group (formally the Defence Science and Technology Organisation [DSTO])

EASA European Aviation Safety Agency

FAA

FLE

FTE

Federal Aviation Administration

Fixed Leading Edge

Fixed Trailing Edge

FEM

FSFT

IV&V

Finite Element Model

Full-scale Wing Fatigue Test

Independent Verification and Validation

MPC

NASTRAN

Multi Point Constraint

NASA Structural Analysis (software)

OEM Original Equipment Manufacturer

RAAF Royal Australian Air Force

RBE Rigid Body Element

SI System International

SSP Structural Substantiation Programme


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1. Introduction

A global Finite Element Model (FEM) of the C-27J airframe is a supplementary tool in support of current and future Royal Australian Air Force (RAAF) C-27J structural integrity management activities.

Key potential uses of the model include the following:

• fatigue test interpretation

• guidance for strain gauge placement on aircraft test articles

• strain, displacement and reaction force predictions for experimental test planning

• estimation of Bearing/Bypass ratios for fastener locations

• estimation of load or field stress inputs to fatigue and damage tolerance analyses

• stress engineering support for the creation of local spectra based on local loads

• supporting aircraft repairs in accordance with approved maintenance manual or Aircraft Structural Integrity Management Plan (ASIMP)

• derivation of internal loads for detailed stress analysis using sub-models.

The purpose of this user manual is to provide a detailed description of the current DST C-27J NASTRAN FEM and at the time of writing, this manual referred version 3.0 of the FEM [1]. It provides enough detail to allow the user to understand how it is organised, what the limitations are and how to incorporate changes. The global FEM comprises a set of sub-models, each of which represents an individual sub-structure of the aircraft. In this document, each sub-model is described separately with figures showing groupings of different sub-model regions and element types. The items that are relevant to the global model are described first, such as the numbering scheme, material properties, and coordinate systems. This is followed by detailed descriptions of the individual sub-models.


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2. Background

DST development of the C-27J Global FEM (based upon the as-delivered OEM FEM) was performed in response to DST task 07/384 as detailed in [2], on the basis that formal acceptance of the Alenia C-27J FEM by the RAAF requires independent verification and validation (IV&V) of the FEM. The rationale for the tasking was focused on developing a model that could be utilised by the RAAF to support the C-27J Aircraft Structural Integrity Program (ASIP). Additionally, the model will be utilised to support DST- activities aimed at interpretation of results of the Full-scale Fatigue Test (FSFT) of the wing.

In 2013, DST Group completed a preliminary review of the C-27J FEM and recommended several areas of model improvement, including specific C-27J FEM verification and validation activities, as per [3]. DST Group was subsequently requested via [4] to complete verification activities for the C-27J FEM. In order to make the FEM more user-friendly and for it to pass the final verification checks, a number of changes to the model were required.

The software tools MSC Patran 2013 (64 bit) and MSC NASTRAN 2013.1.0 for Windows were used to perform this work. Henceforth, these will be referred to simply as “Patran” and “NASTRAN”.


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3. Model Enhancements and Quality Checks

Numerous enhancements were made while converting the original OEM FEM to the v3.0 C-27J Global FEM, in order to improve the model quality, accuracy and ease of use. The points below list the principal areas of improvement. A full account, including a record of FEM version changes, is found in [6]. Model quality checks are also found within [6].

• Separation of the FEM into individual sub-models, each of which represents a separate aircraft substructure (e.g. forward fuselage, centre wing, empennage, etc.)

• Development of interface connection between sub-models, comprising Rigid Body Elements (RBE) using NASTRAN type RBE2, to assemble sub-models into an integrated global FEM

• Creation of a unique numbering range for each sub-model

• Reduction of the number of material entries, consistent with global FEM requirements and consolidate into a separate single NASTRAN input file

• Consolidate coordinate systems definition into a separate single NASTRAN input file

• Replace of all Multipoint Constraint entries (MPC) with RBE2 elements

• Correction of element types - membrane and shear elements converted to shell elements; rod elements (used to represent flanges) converted to bar elements

• Since this is not a dynamic model, the mass does not need to be representative. Therefore, all the densities have been set to the correct value for each material, whilst non-structural mass has not been included

• Adoption of a unified unit system. Legacy model used multiple Engineering units including daN, kgf, etc. Version 3.0 uses N, mm, deg. C and derived pressure units of MPa

• Use of English file names and file headers throughout

• Correction of model and loads orientation such that z-axis is directed vertically up, consistent with DST Group norms

• Incorporation of design beam sections consistent with design geometry, by way of NASTRAN PBMSECT idealisation (after converting bar elements to beam elements).


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4. How to Use This FEM

4.1 Fundamental Principles

Under no circumstances should users treat the outputs of the C-27J Global FEM in isolation, or as a single authoritative reference for aircraft stress data. The C-27J Global FEM is a supplementary tool to be used in conjunction with the relevant aircraft certification stress dossiers, the relevant test reports, and/or other certification documentation as appropriate to the problem at hand.

4.2 Limitations

The C-27J Global FEM still has some limitations which have a bearing on how this model can currently be reliably used. These are listed and fully described in Sections 5, 8 and 11 of this document. All users must carefully assess the model to determine whether faithful results can be expected in their particular area of interest. Load and constraint modifications will likely be required for aft fuselage, landing gear, and many other fuselage areas.

4.3 Modifying the FEM

This FEM is divided into substructure models which can be run independently of the global model. The independent sub-models and the global arrangement are shown in Figure 1. Changes made to the sub-models should preferably adhere to the defined numbering system as described in Section 7. Although all entities in the sub-models can be completely renumbered, care should be taken when renumbering the interface nodes because these are referenced by rigid body elements (RBE2) that join the sub-models together. Any modifications should be recorded in the sub-model file headers to ensure that version traceability and change control is maintained. After editing, model quality checks should be performed on each affected sub-model and on the global model to ensure that no errors have been introduced. Some of the reasons for editing the sub-models include:

• renumbering of elements, nodes, and properties

• local mesh refinement or the introduction of model features, where more accurate outputs are required

• incorporating change to the structure, such as repair schemes

• changing mesh or incorporating other types of elements in-order to extract different types of information

• model modification to aid the study of additional load cases

• modification to loading and boundary conditions.


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Figure 1: Assembled global FEM (below), and sub-model identification (above)

Version 3.0 of the C-27J Global FEM is configured such that the bulk data files and loading files are independently legible as text files. Each file is organised similarly, with file headers and text comments that allow relative ease of reading. Accordingly, the model should not be exported in its entirety from Patran, as the file structure and model organisation will be lost as a result. It is therefore recommended that Patran session files are created, containing sub-structure groups and visualisation parameters, to allow efficient interaction with the model.


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4.4 Support to Aircraft Structural Integrity

Correct fatigue life prediction depends on accurate determination of local stresses. The C-27J Global FEM is a valuable tool to support development of tailored stress and fatigue analysis solutions, by providing accurate remote field stress data. This FEM can therefore support analyses aimed at increasing inspection intervals or life extension of ageing aircraft.

4.5 Use as a Loads or Gross Remote Stress Model

The C-27J Global FEM is configured in such a way that for the provided load cases, the member forces and moments can be used as inputs for subsequent calculations (for example, classical calculations or damage tolerance analysis)1. Similarly, gross remote or ‘field’ stresses can be taken from the model; however care must be taken, as factoring based on design/FEM element thicknesses may be required2

4.6 Use as a Stress Model

The C-27J Global FEM is configured to provide gross global behaviour of the aircraft under various loading conditions. It is not intended to provide accurate detailed local stress information. In order to develop local stress information, it is recommended that sub-modelling techniques are used, or if appropriate, local mesh refinement. In the event that the latter is taken as the chosen path, care must be taken to ensure that geometry, material, meshing and loading parameters are appropriate to the case at hand.

4.7 Support to Detailed Stress Analysis

The C-27J Global FEM can support detailed stress, loads and displacement modelling as discussed below.

4.7.1 Local refinement using 3D solid elements

This FE model will permit the attachment of solid elements to the existing shell elements and thus allow detailed stressing of 3D parts. In developing a hybrid model such as this, care must be taken at element interfaces. Alternatively, super elements and sub-modelling

1 Discretion is required in selecting sites that will provide representative loads and/or stresses from this model. See Sections 4.1, 4.2, 5 and 11 of this document for further detail. 2 It is not unusual for a global FEM to adopt geometry that differs from design. This is in order to achieve representative global stiffness characteristics in the absence of the finer design detail. Therefore, gross remote field stresses may require factoring in order to produce the equivalent ‘design’ stress levels. For example, panels with cut-outs may not be represented with cut-outs included in the FEM. Thus, it may be necessary to factor the FEM stress output based on the difference in panel thickness, to produce the equivalent ‘design’ stress. Such factoring must be carefully developed on a case-by-case basis.


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techniques can be used to drive loads into 3D sub-models and thus allow detailed stress analysis of 3D parts.

4.7.2 Derivation of interface loads at sub-model boundaries

The arrangement of the global FEM into constituent sub-models allows interface loads to be easily extracted for all load cases. Care must be taken however, to ensure that the boundary conditions for the particular load case in question are sufficiently representative for the attachment being considered. The extraction of boundary and interface loads allows classical or other detailed stress calculations to be carried out where necessary.

4.7.3 Aid aircraft development

The correlation of load cases, aircraft geometry and internal loads offers the potential to allow aircraft development on a more numerically informed basis.

4.7.4 Cross-check stress transfer functions by deriving scale factors

The global stiffness matrix from the global FEM can be used to cross-check transfer functions inasmuch as checking that stiffness is representative. Noting that damping is minimal in the airframe, it follows that vibration transfer functions (used in some fatigue analyses) can be checked with reasonable accuracy via the global FEM.

4.7.5 Damage tolerance analysis

Damage tolerance analysis is an important aspect of structural analysis that involves introducing defects into the FE model, and then analysing the residual strength of the surrounding structure. This global FEM will support the introduction of structural defects by removing parts, partially or completely severing parts, or removing connections, etcetera.

4.7.6 Determine stress transfer functions

Cognizant of the fact that the global FEM is a linear elastic model, it follows that it can be used to produce load and stress relationships anywhere on the airframe. It follows therefore that load to stress relationships can be developed (equally, displacement to stress relationships can be developed). Such relations can be used for detailed stress analyses, for example, static stress or fatigue analyses.


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4.8 Support to Aircraft or Aircraft Component Testing

The C-27J Global FEM can support aircraft or component testing as discussed below.

4.8.1 Strain, stress, load and displacement predictions

The C-27J Global FEM can be modified to accommodate the loading and displacement constraints that match test configurations. Therefore it can used to predict stress, strain, internal loads and displacements under test conditions.

4.8.2 Guidance for strain gauge placement on test articles

The C-27J Global FEM offers a large number of load cases and also the capacity for these load cases to be tailored to DST Group unique requirements. It follows therefore that running loads, stresses and strains can be taken from any number of load cases and used to inform the placement of strain gauges.

4.8.3 Validation of test representativeness

The C-27J Global FEM can be loaded to reflect either test configurations or flight configurations. Therefore it can be used to compare between test and flight arrangements.

4.8.4 Virtual ground calibration

Ground calibration can be expressed in a number of forms, for example load versus displacement, load versus wing curvature, load versus stress, or any other combination of the above. With the appropriate boundary conditions applied to the model, it follows that the C-27J Global FEM offers the possibility to carry out a virtual ground calibration.


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5. Aircraft Configuration and Modelling Philosophy

5.1 Wing

5.1.1 Leading and Trailing Edge components

The model idealisation of the wing only captures the central portion of the wing box cross-section. This is to say, the model contains no Fixed Leading Edge (FLE), Fixed Trailing Edge (FTE), or moveables (e.g. ailerons or flaps). While the C-27J Global FEM is configured to provide representative behaviour of the main structural components of the wing, care must be taken in extracting FEM outputs that could be influenced by the absence of the above structure. Additionally, at FEM version 3.0, the model has inner to outer wing connection via the spars only (i.e. no covers connection). This affects covers and spars stresses in the region of the join when subject to wing bending and torsion load cases - a point to be addressed in future versions of this FEM.

5.1.2 Ribs

Aerodynamic loads are not introduced into the model by pressure fields, but by way of an array of interpolation elements (NASTRAN type RBE3) that act upon the periphery of the wing ribs. Such elements are chosen specifically to drive loads into the wing structure whilst allowing the ribs themselves to deform (i.e. the RBE3 elements do not add artificial stiffness). This is shown in Figure 2. The obvious difference in loading compared to pressure fields means that care must be taken when examining rib loads. This is especially the case in the outer portion of the wing, where there is a pronounced difference between the model idealisation (at the rib) and the true loading characteristics (pressure fields applied to covers). For example, extraction of vertical stiffener and/or rib foot loads from the outer-wing portion of this model would be inappropriate due to the applied load (via the rib RBE3’s) artificially influencing the built-up load in the structure. The sum of FE loads and classically calculated loads3 may provide reasonable results for the inner wing, however this does not apply for the outer wing, where classical Brazier loading calculations would be required. Consistent with most global FEM representations, rib cut-outs are not included in the model. Stress detail for cut-out locations would typically be calculated by factoring the relevant panel stresses.

Figure 2: Wing loading idealisation, showing close-up view of rib loading RBE3 strategy 3 Typical classical calculations would include pull-off calculations, to account for aerodynamic loads and fuel pressure loads that act upon covers.


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5.1.3 Covers

As per Figure 3, cover detail shows stringers, skins, mid-spar and rib attachments. Due to the nature of wing load introduction and the absence of pressure fields on the wing, covers pull-off loading (due to fuel pressurisation and air pressure) and the associated rib-foot loading is not represented in this model. Further, consistent with typical global FEM representations, cut-outs are not included in the model This is shown in Figure 4. Stress detail for cut-out locations would typically be calculated by detailed modelling or factoring based on the panel stresses from this model, and accounting for the geometric differences.

Figure 3: Covers showing stringer idealisation. Wing Outer Cover, Lower, Portside.

Figure 4: Portside wing showing the absence of access holes. Loading elements (RBE3) in pink.

5.2 Forward Fuselage

Forward fuselage loading does not correctly capture pressure loading because due to the legacy modelling constructs, there is no load on the windshield frame and other fuselage openings to account for the absent pressure areas. See Figure 5. The C-27 J Global FEM does not contain representations of the radome (nose of the aircraft) nor any landing gear. Therefore, care must be taken in examining load cases and structure influenced in this regard.


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Figure 8: Cut-away view of empennage loading idealisation. Note RBE’s applied to the outer

periphery of the stabilizer attachment frames.

5.5 Empennage

Unlike the wing, the Horizontal and Vertical Stabilizers do not have loads applied through their ribs. Furthermore, the provided loads files do not include any applied pressure loads. Therefore, the study of any empennage structure must establish and apply the appropriate loads to this model, and understand the idealisation limitations on what can be reasonably extracted from the model. The Alenia loading strategy can be seen in Figure 8.

5.6 Landing Gear Attachments

As per Figure 7, Landing Gear Attachments are not loaded in this FEM. Because the loading RBE’s of this model do not address load input due to the landing gear, the study of structure local to the landing gear attachments when subject to ground cases will likely require load introduction and possibly model modification beforehand.

5.7 Aircraft Configuration – Summary

Several significant simplifications and model limitations have been identified in the C-27J Global FEM at version 3.0. These are wholly the result of the legacy modelling approach. In summary, these are as follows:

• Wing loading idealisation does not contain fuel or aero pressure fields. This influences ribs and covers. To address this demands supplementary analysis.

• Outer wing brazier loads cannot be expected to be correct, due to the strategy of loading the structure via the ribs. To address this would require supplementary analysis.


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• Load cases where the wing FLE, FTE and moveables are relevant require careful consideration, as this structure is not represented in this model.

• Forward fuselage lacks pressure loading at the windshield frame and other fuselage openings. This points to limited useable data expected from forward fuselage when subject to pressure load cases, especially at the location of openings, unless further model development is performed.

• Centre fuselage loading approach ignores pressure loads about door openings, and adopts a simplified loading approach via fuselage frames, which points to limited useable data expected from centre fuselage when subject to pressure load cases, unless further model development is performed.

• Aft fuselage loading approach points to very little useful data expected from empennage region as a whole, including Horizontal and Vertical Stabiliser attachment frames, unless further model development is performed.

• Landing gear input loading approach is not representative thus structure local to the attachments cannot be expected to show accurate results when subject to ground or landing load cases.

On considering the above, it is evident that the extent to which this model is valid for the provided load cases is restricted. It is therefore necessary that the end-users of this model take adequate steps to ensure that the geometric, loading and constraint idealisations are appropriate for the load cases under consideration.


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6. Load Cases

The below Tables 1 through Table 4 present the load cases as provided by Alenia Aeronautica, as a constituent of the global FEM delivery as per [3]. As discussed previously, the extent to which useful data can be extracted from the model for these cases requires individual consideration and model scrutiny.

Table 1: Static Ultimate load cases for use with C-27J v3.0 Global FEM

File Subcase Description run_static_fusol_JCA_ult 10 PRESSURE 6.2 psi run_static_fusol_JCA_ult 11 LEFT LATERAL MAN. ENVELOPE PRESSURISED run_static_fusol_JCA_ult 4201 SPIN UP FX=80% FZMAX NZ=2 V=1.25*VL2 M=M2 run_static_fusol_JCA_ult 4202 SPRING BACK FX=65% FXMAX NZ=2 V=1.25VL2 M=M2 run_static_fusol_JCA_ult 4203 SPIN UP FX=80% FZMAX NZ=2 V=1.25VL2 M=M2 run_static_fusol_JCA_ult 4204 SPRING BACK FX=65% FXMAX NZ=2 V=1.25VL2 M=M2 run_static_fusol_JCA_ult 4205 SPIN UP FX=80% FZMAX N=2 V=1.25VL2 M=M2

LOAD 4205 run_static_fusol_JCA_ult 4206 SPRING BACK FX=65% FXMAX NZ=2 V=1.25*VL2 M=M2 run_static_fusol_JCA_ult 4207 SPIN UP FX=80% FZMAX NZ=2 V=1.25*VL2 M=M2 run_static_fusol_JCA_ult 4208 SPRING BACK FX=65% FxMAX NZ=2 V=1.25*VL2 M=M2 run_static_fusol_JCA_ult 4299 LEFT LATERAL MAN. ENVELOPE run_static_fusol_JCA_ult 4301 C05ASC1261-4 run_static_fusol_JCA_ult 4302 C4BCSS0363 run_static_fusol_JCA_ult 4303 C05ASS0543 run_static_fusol_JCA_ult 4304 C4CCSD5381 run_static_fusol_JCA_ult 4305 C05CSS0552 run_static_fusol_JCA_ult 4306 C7AASU0895-2 run_static_fusol_JCA_ult 4307 C7ACSD5273 L=18.3 m t=0.2502 s run_static_fusol_JCA_ult 4308 C4CCSD5189 L=18.3 m t=0.2540 s run_static_fusol_JCA_ult 5301 C05ASC1261-4 PRESSURISED run_static_fusol_JCA_ult 5302 C4BCSS0363 PRESSURISED run_static_fusol_JCA_ult 5303 C05ASS0543 PRESSURISED run_static_fusol_JCA_ult 5304 C4CCSD5381 PRESSURISED run_static_fusol_JCA_ult 5305 C05CSS0552 PRESSURISED run_static_fusol_JCA_ult 5306 C7AASU0895-2 PRESSURISED run_static_fusol_JCA_ult 5307 C7ACSD5273 L=18.3 m t=0.2502 s PRESSURISED run_static_fusol_JCA_ult 5308 C4CCSD5189 L=18.3 m t=0.2540 s PRESSURISED


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Table 2: Static Limit load cases for use with C-27J v3.0 Global FEM

File Subcase Description run_static_wing_full_3D_limit 3032 JACKING CASE run_static_wing_full_3D_limit 3208 STEADY MAN. - C 3B C S S 0066 run_static_wing_full_3D_limit 3209 STEADY MAN. - C 4B C S S 0336 run_static_wing_full_3D_limit 3301 SYM. MAN. at VD, M6, NZ=2.5, T=0., C06DSS0671 run_static_wing_full_3D_limit 3302 SYM. MAN. at VD, M7, NZ=-1., T=0., C07DSS0726 run_static_wing_full_3D_limit 3303 UP GUST - Z=18400ft - M=0.55 - Vc - M3-B run_static_wing_full_3D_limit 3304 VERTICAL GUST at VC, M4B, M=0.55, z=18400. run_static_wing_full_3D_limit 3305 VERTICAL GUST at VC, M4D, M=0.55, z=18400. run_static_wing_full_3D_limit 3306 DISCRETE GUST C4DCSD5201 - L=75.3 - t=0.366 run_static_wing_full_3D_limit 3307 DISCRETE GUST C4DCSD5201 - L=75.3 - t=0.3884 run_static_wing_full_3D_limit 3308 DISCRETE GUST C4BCSD5177 – L=31.3 – t=0.2838 run_static_wing_full_3D_limit 3309 DISCRETE GUST – C4DCSD5201 – L=0.3 – t=0.3996 run_static_wing_full_3D_limit 3310 DISCRETE GUST – C4DCSD5201 – L=0.3 – t=0.4033 run_static_wing_full_3D_limit 3311 DISCRETE GUST– C4DCSD5201 – L=57.3 – t=0.3697 run_static_wing_full_3D_limit 3312 DISCRETE GUST– C4DCSD5201 – L=57.3 – t=0.3735 run_static_wing_full_3D_limit 3313 DISCRETE GUST– C4DCSD5201 – L=31.3 – t=0.2838 run_static_wing_full_3D_limit 3314 DISCRETE GUST C4DCSD5213 - L=31.3 - t=0.4594 run_static_wing_full_3D_limit 3315 DYNAMIC LANDING – L8ANSI2178(SB) run_static_wing_full_3D_limit 13401 UP GUST 1855 – PRATT – C4BBSG1855 run_static_wing_full_3D_limit 13402 UP GUST 1695 – PRATT – C3BBSG1695 run_static_wing_full_3D_limit 13403 UP GUST 1919 – PRATT – C4DBSG1919 run_static_wing_full_3D_limit 13405 DISCRETE GUST – C4DCSD5201 – L=57.3 – t=0.366

Table 3: Fatigue ‘JCA’ load cases for use with C-27J v3.0 Global FEM

File Subcase Description run_fatigue_JCA 501 FATIGUE UNLOADED CONDITION run_fatigue_JCA 502 FATIGUE TAXI LEFT GROUND TURN CONDITION run_fatigue_JCA 503 FATIGUE TAXI RIGHT GROUND TURN CONDITION run_fatigue_JCA 504 FATIGUE TAXI 1.3g CONDITION run_fatigue_JCA 505 FATIGUE TAXI 0.7g CONDITION run_fatigue_JCA 506 FATIGUE ROTATION CONDITION run_fatigue_JCA 507 FATIGUE CRUISE 1g+20ft RAFF. VERT. CONDITION run_fatigue_JCA 508 FATIGUE CRUISE 1g-20ft RAFF. VERT. CONDITION run_fatigue_JCA 509 FATIGUE CRUISE 1g+23ft RAFF. LAT. CONDITION run_fatigue_JCA 510 FATIGUE CRUISE 1g+23ft RAFF. LAT. CONDITION run_fatigue_JCA 511 FATIGUE CRUISE 1g+0.4 nz MAN. CONDITION run_fatigue_JCA 512 FATIGUE CRUISE 1g-0.4 nz MAN. CONDITION run_fatigue_JCA 513 FATIGUE YAW MANOUVRE 1g+7 deg. CONDITION run_fatigue_JCA 514 FATIGUE YAW MANOUVRE 1g-7 deg. CONDITION run_fatigue_JCA 515 FATIGUE FLARE CONDITION run_fatigue_JCA 516 FATIGUE TOUCH DOWN 1g+4.5ft/sec CONDITION run_fatigue_JCA 517 FATIGUE ENGINE RUN UP CONDITION run_fatigue_JCA 518 FATIGUE REVERSE CONDITION


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Table 4: Fatigue ‘wing full’ load cases for use with C-27J v3.0 Global FEM

File Subcase Description run_fatigue_wing_full_3D 501 FATIGUE UNLOADED CONDITION run_fatigue_wing_full_3D 502 FATIGUE TAXI LEFT GROUND TURN CONDITION run_fatigue_wing_full_3D 503 FATIGUE TAXI RIGHT GROUND TURN CONDITION run_fatigue_wing_full_3D 504 FATIGUE TAXI 1.3g CONDITION run_fatigue_wing_full_3D 505 FATIGUE TAXI 0.7g CONDITION run_fatigue_wing_full_3D 506 FATIGUE ROTATION CONDITION run_fatigue_wing_full_3D 507 FATIGUE CRUISE 1g+20ft RAFF. VERT. CONDITION run_fatigue_wing_full_3D 508 FATIGUE CRUISE 1g-20ft RAFF. VERT. CONDITION run_fatigue_wing_full_3D 509 FATIGUE CRUISE 1g+23ft RAFF. LAT. CONDITION run_fatigue_wing_full_3D 510 FATIGUE CRUISE 1g+23ft RAFF. LAT. CONDITION run_fatigue_wing_full_3D 511 FATIGUE CRUISE 1g+0.4 nz MAN. CONDITION run_fatigue_wing_full_3D 512 FATIGUE CRUISE 1g-0.4 nz MAN. CONDITION run_fatigue_wing_full_3D 513 FATIGUE YAW MANOUVRE 1g+7 deg. CONDITION run_fatigue_wing_full_3D 514 FATIGUE YAW MANOUVRE 1g-7 deg. CONDITION run_fatigue_wing_full_3D 515 FATIGUE FLARE CONDITION run_fatigue_wing_full_3D 516 FATIGUE TOUCH DOWN 1g+4.5ft/sec CONDITION run_fatigue_wing_full_3D 517 FATIGUE ENGINE RUN UP CONDITION run_fatigue_wing_full_3D 518 FATIGUE REVERSE CONDITION


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7. FEM Numbering System

The global FEM comprises 11 major substructure models as shown in Figure 9. The advantage of organising the global FEM in this manner is that changes made to the model can be isolated and evaluated on the sub-model level before incorporation into the global model. Additionally, it allows analysis of different sub-model versions by simply swapping one for the other.

A numbering scheme has been developed in which clearly defined numbering ranges have been defined for each of the aircraft substructures represented in the sub-models. All of the elements and nodes contained in each substructure follow this numbering scheme. The allocated numbering ranges are shown in Table 5, and the utilised numbering ranges at version 3.0 of the FEM are shown in Table 6. By examining these ranges, the allowable ranges for sub-model growth or modification can be deduced. Note that the v3.0 FEM uses the NASTRAN beam element section type “PBMSECT”, whereupon sections are identified as a set of points about a section profile. It is crucial that the identifiers for these points (point ID’s) do not conflict with any of the FE model node identifiers (grid ID’s).

Figure 9: C-27J Global FEM separated into sub-models


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Table 5: Sub-model node and element and MPC allocation number ranges (ID ranges)

C-27J FEM Numbering Range Elements Nodes* MPCs

Component Description From To

Fuselage Forward 1,000,000 1,999,999 X X X Fuselage Centre 2,000,000 2,999,999 X X X Fuselage Aft 3,000,000 3,499,999 X X X Elevator Linkage Port 3,330,000 3,339,999 X X X Elevator Linkage Starboard 3,340,000 3,349,999 X X X Cargo Door 3,500,000 3,699,999 X X X Cargo Ramp 3,700,000 3,899,999 X X X Wing 4,000,000 4,999,999 X X X Wing Centre 4,000,000 4,499,999 X X X Wing Outer Port 4,500,000 4,699,999 X X X Wing Outer Starboard 4,700,000 4,899,999 X X X Engines 5,000,000 5,999,999 X X X Engine 1 5,000,000 5,099,999 X X X Engine 2 5,100,000 5,199,999 X X X Empennage 6,000,000 6,999,999 X X X Vertical Stabiliser 6,000,000 6,399,999 X X X Horizontal Stabiliser 6,500,000 6,599,999 X X X Horizontal Stabiliser 6,600,000 6,699,999 X X X Elevator Port 6,700,000 6,789,999 X X X Elevator Starboard 6,800,000 6,889,999 X X X Mass and Loads 7,000,000 7,999,999 X X X Mass Cargo 7,000,000 7,099,999 X X X Mass Fuel 7,100,000 7,199,999 X X X Load RBE3’s and Grids 7,500,000 7,999,999 X X Interface Connections (RBE2 or CBUSH) 8,000,000 8,999,999 X X X

*Point ID’s for PBMSECT property cards included


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Table 6: Sub-model node and element identification number utilisation (ID utilisation)

Location Type: CBAR CQUAD4 CTRIA3 GRID PBAR PSHELL PBEAM CBEAM RBE3 RBE2 POINT PBMSECT Aft Fuse Count: 2793 1633 544 1711 223 90 0 0 0 0 0 0

Max: 3344000 3051186 3280866 3344000 3343997 3013585 0 0 0 0 0 0 Min: 3000039 3040001 3031270 3000835 3010023 3011050 0 0 0 0 0 0

Cargo Door Count: 862 515 12 420 51 26 1 24 0 0 0 0 Max: 3582406 3582194 3582106 3562026 3610116 3514011 3508002 3581429 0 0 0 0 Min: 3581000 3581462 3551002 3506001 3508000 3508003 3508002 3581127 0 0 0 0

Wing Outer Starboard

Count: 635 852 87 880 117 77 0 500 0 0 0 0 Max: 4733853 4733909 4733910 4733045 4836402 4736077 0 4723745 0 0 0 0 Min: 4714851 4701111 4701602 4714001 4736101 4736001 0 4702701 0 0 0 0

Wing Outer Port

Count: 635 852 87 880 117 77 0 500 0 0 0 0 Max: 4533853 4533909 4533910 4533045 4630402 4530077 0 4523745 0 0 0 0 Min: 4514851 4501111 4501602 4514001 4530101 4530001 0 4502701 0 0 0 0

Centre Wing

Count: 1353 1408 122 1452 91 61 0 812 0 8 0 0 Max: 4424249 4434174 4434175 4414074 4138890 4029933 0 4409437 0 4401261 0 0 Min: 4300001 4301401 4301415 4001836 4024010 4025310 0 4307001 0 4300861 0 0

Vertical Stabiliser Count: 1348 896 66 855 100 11 0 0 0 5 0 0 Max: 6052004 6051157 6051350 6344088 6146509 6046004 0 0 0 6053004 0 0 Min: 6000011 6010001 6010302 6000011 6044001 6042001 0 0 0 6000007 0 0

Starboard Elevator Count: 261 231 5 168 25 12 0 0 0 0 0 0 Max: 6853499 6853231 6853236 6853168 6859004 6857011 0 0 0 0 0 0 Min: 6853237 6853001 6853232 6853001 6858000 6857000 0 0 0 0 0 0

Horizontal Stabiliser Starboard

Count: 412 472 22 337 26 10 0 239 0 0 0 0 Max: 6611147 6611128 6611121 6610343 6656021 6654010 0 6611099 0 0 0 0 Min: 6610432 6610001 6610086 6610001 6655001 6654001 0 6610353 0 0 0 0


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Table 6: (cont’d)

Location Type: CBAR CQUAD4 CTRIA3 GRID PBAR PSHELL PBEAM CBEAM RBE3 RBE2 POINT PBMSECT Port Elevator Count: 261 231 5 168 25 12 0 0 0 0 0 0

Max: 6753499 6753231 6753236 6753168 6753004 6751011 0 0 0 0 0 0 Min: 6753237 6753001 6753232 6753001 6752000 6751000 0 0 0 0 0 0

Horizontal Stabiliser Port

Count: 412 472 22 337 26 10 0 239 0 0 0 0 Max: 6511147 6511128 6511121 6510343 6550021 6548010 0 6511099 0 0 0 0 Min: 6510432 6510001 6510086 6510001 6549001 6548001 0 6510353 0 0 0 0

Centre Fuselage Count: 3671 1926 367 1998 128 34 0 0 0 0 0 0 Max: 2954083 2320128 2316048 2400826 2950001 2013516 0 0 0 0 0 0 Min: 2026076 2027889 2026231 2000800 2002000 2000001 0 0 0 0 0 0

Beam cross- section Outer Wing Starboard

Count: 0 0 0 0 0 0 0 0 0 0 292 26 Max: 0 0 0 0 0 0 0 0 0 0 4733337 4736327 Min: 0 0 0 0 0 0 0 0 0 0 4733046 4736301

Beam cross-section Outer Wing

Port

Count: 0 0 0 0 0 0 0 0 0 0 292 26 Max: 0 0 0 0 0 0 0 0 0 0 4533337 4530327 Min: 0 0 0 0 0 0 0 0 0 0 4533046 4530301

Beam cross-section Centre Wing

Count: 0 0 0 0 0 0 0 0 0 0 850 98 Max: 0 0 0 0 0 0 0 0 0 0 4328785 4028889 Min: 0 0 0 0 0 0 0 0 0 0 4327936 4028002

Beam cross-section Horizontal Stabiliser

Starboard

Count: 0 0 0 0 0 0 0 0 0 0 97 8 Max: 0 0 0 0 0 0 0 0 0 0 6620423 6656025 Min: 0 0 0 0 0 0 0 0 0 0 6620327 6655000

Beam cross-section Horizontal Stabiliser

Port

Count: 0 0 0 0 0 0 0 0 0 0 97 8 Max: 0 0 0 0 0 0 0 0 0 0 6520423 6550025 Min: 0 0 0 0 0 0 0 0 0 0 6520327 6549000


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Table 6: (cont’d)

Location Type: CBAR CQUAD4 CTRIA3 GRID PBAR PSHELL PBEAM CBEAM RBE3 RBE2 POINT PBMSECT Interface Connector

Centre Wing & Starboard

Count: 0 0 0 0 0 0 0 0 0 44 0 0 Max: 0 0 0 0 0 0 0 0 0 8447045 0 0 Min: 0 0 0 0 0 0 0 0 0 8447001 0 0

Interface Connector Centre Wing

& Port

Count: 0 0 0 0 0 0 0 0 0 44 0 0 Max: 0 0 0 0 0 0 0 0 0 8445045 0 0 Min: 0 0 0 0 0 0 0 0 0 8445001 0 0

Interface Connector Horizontal Stabiliser

& Elevator Starboard

Count: 0 0 0 3 0 0 0 0 0 6 0 0 Max: 0 0 0 8666803 0 0 0 0 0 8666813 0 0 Min: 0 0 0 8666801 0 0 0 0 0 8666801 0 0

Interface Connector Horizontal Stabiliser

& Elevator Port

Count: 0 0 0 3 0 0 0 0 0 6 0 0 Max: 0 0 0 8656703 0 0 0 0 0 8656713 0 0 Min: 0 0 0 8656701 0 0 0 0 0 8656701 0 0

Interface Connector Forward and Aft

Fuselage

Count: 0 0 0 0 0 0 0 0 0 66 0 0 Max: 0 0 0 0 0 0 0 0 0 8120067 0 0 Min: 0 0 0 0 0 0 0 0 0 8120001 0 0

Interface Connector Centre Fuselage

& Wing

Count: 0 0 0 0 0 0 0 0 0 43 0 0 Max: 0 0 0 0 0 0 0 0 0 8240044 0 0 Min: 0 0 0 0 0 0 0 0 0 8240001 0 0

Interface Connector Centre Fuselage & Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 61 0 0 Max: 0 0 0 0 0 0 0 0 0 8230062 0 0 Min: 0 0 0 0 0 0 0 0 0 8230001 0 0

Interface Connector Centre Fuselage

& Ramp

Count: 0 0 0 0 0 0 0 0 0 4 0 0 Max: 0 0 0 0 0 0 0 0 0 8237004 0 0 Min: 0 0 0 0 0 0 0 0 0 8237001 0 0


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Table 6: (cont’d)

Location Type: CBAR CQUAD4 CTRIA3 GRID PBAR PSHELL PBEAM CBEAM RBE3 RBE2 POINT PBMSECT Interface

Connector Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 5 0 0 Max: 0 0 0 0 0 0 0 0 0 8360006 0 0 Min: 0 0 0 0 0 0 0 0 0 8360001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 3 0 0 Max: 0 0 0 0 0 0 0 0 0 8366004 0 0 Min: 0 0 0 0 0 0 0 0 0 8366001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 3 0 0 Max: 0 0 0 0 0 0 0 0 0 8365004 0 0 Min: 0 0 0 0 0 0 0 0 0 8365001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 0 0 0 Max: 0 0 0 0 0 0 0 0 0 8368001 0 0 Min: 0 0 0 0 0 0 0 0 0 8368001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 0 0 0 Max: 0 0 0 0 0 0 0 0 0 8367001 0 0 Min: 0 0 0 0 0 0 0 0 0 8367001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 9 0 0 Max: 0 0 0 0 0 0 0 0 0 8337010 0 0 Min: 0 0 0 0 0 0 0 0 0 8337001 0 0

Interface Connector

Aft Fuselage

Count: 0 0 0 0 0 0 0 0 0 10 0 0 Max: 0 0 0 0 0 0 0 0 0 8335012 0 0 Min: 0 0 0 0 0 0 0 0 0 8335003 0 0

Loading RBE Fuselage

Count: 0 0 0 42 0 0 0 0 42 0 0 0 Max: 0 0 0 7601450 0 0 0 0 7526311 0 0 0 Min: 0 0 0 7601000 0 0 0 0 7501780 0 0 0


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Table 6: (cont’d)

Location Type: CBAR CQUAD4 CTRIA3 GRID PBAR PSHELL PBEAM CBEAM RBE3 RBE2 POINT PBMSECT Loading RBE Centre Wing

Count: 0 0 0 136 0 0 0 0 136 0 0 0 Max: 0 0 0 7700133 0 0 0 0 7501633 0 0 0 Min: 0 0 0 7600000 0 0 0 0 7500500 0 0 0

Forward Fuselage Count: 2112 1010 238 1193 229 18 0 0 0 0 0 0 Max: 1280516 1090052 1090040 1999991 1003002 1000665 0 0 0 0 0 0 Min: 1001500 1002198 1002208 1000001 1000201 1000266 0 0 0 0 0 0

Interface Connector Aft Fuselage

& Door Hinge

Count: 0 0 0 0 0 0 0 0 0 2 0 0 Max: 0 0 0 0 0 0 0 0 0 8335002 0 0 Min: 0 0 0 0 0 0 0 0 0 8335001 0 0


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8. Material Properties and Units Of Measure

The model comprises 3 isotropic materials as shown in Figure 10. Whilst this may not reflect design material allocations exactly, the stiffness properties are considered sufficiently similar to allow accurate data to be extracted from this model. The unit system used in this model is SI metric units throughout, namely N, mm, deg. C. Stress units and other derived units are deduced from these base units (e.g. stress in N/mm or MPa.)

Figure 10: Material properties


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9. Coordinate Systems

There are a total of 12 local coordinate systems used in the model. Figure 11 through Figure 17 shows the location and orientation of the respective coordinate systems for version 3.0 of this FEM. All coordinate systems are Cartesian and some, as part of the legacy of this FEM, are coincident with one another. All local coordinate systems are defined in the bulk data file “coordinate_systems.bulk” which is shown in Figure 18.

Figure 11: Local coordinate systems as per forward fuselage

Figure 12: Local coordinate. systems as per centre fuselage


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Figure 13: Local coordinate systems as per wing; top and bottom highlights are of lower cover only

Figure 14: Local coordinate systems as per aft fuselage


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Figure 15: Local coordinate systems as per empennage

Figure 16: Local coordinate systems as per cargo ramp (note coincident coordinate systems)


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Figure 17: Local coordinate systems – cargo door (note coincident coordinate systems)


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Figure 18: Local coordinate systems as per bulk data file “coordinate_systems”


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10. Structural Idealisation and Sub-Model Descriptions

In this section, the properties of the sub-models are described. Specifically, Figure 19 through Figure 58 show the FEM idealisation as compared to assembly illustrations (i.e. as per [5]), the materials used within the FEM, and the application of shell and beam elements throughout the structure. Where structural depictions as per [5] are presented, annotations are to be ignored, or if potentially useful, referred back to the source document.

10.1 Forward Fuselage

10.1.1 General arrangement

Figure 19: Forward fuselage, FEM isometric view and corresponding illustration as per [4]


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10.1.2 Shell elements

Figure 20: Forward fuselage, shell elements (section view, starboard, to show interior elements)


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10.1.3 Bar elements

Figure 21: C-27J forward fuselage, bar elements (no beam elements. All of material 7075)


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10.2 Centre Fuselage


Figure 23: C-27J Centre fuselage FEM isometric view and as per [5]


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Figure 24: C-27J Centre Fuselage shell elements (isometric view)

10.2.3 Beam elements

Figure 25: C-27J Centre fuselage beam elements (isometric view)


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10.2.4 Material allocation

Figure 26: C-27J centre fuselage material designations. Skin: 2024, interior: 7075.


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10.3 Aft Fuselage


Figure 27: C-27J Aft fuselage FEM (upper) isometric view and as per [5] (lower)


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Figure 28: C-27J Aft Fuselage shell elements (isometric view)


Figure 29: C-27J Aft fuselage beam elements (isometric view)


Figure 30: C-27J Aft fuselage material designations. Skin: 2024, remaining structure: 7075


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10.4 Centre Wing


Figure 31: C-27J Centre wing box FEM isometric view (upper) and as per [5] (lower)


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Figure 32: C-27J Centre wing box shell elements (isometric view)


Figure 33: C-27J Aft fuselage beam elements (isometric view)


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Figure 34: C-27J Centre wing box material designations. Top Cover (7075) removed for detailed

view, showing lower skin (2024).

10.5 Outer Wing (starboard side shown, port side by similarity)


Figure 35: C-27J Outer Wing FEM isometric view (upper) and as per [5] (lower)


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Figure 36: C-27J Outer wing shell elements (isometric view)


Figure 37: C-27J Outer wing beam elements (isometric view)


Figure 38: C-27J Outer Wing, Inverted (lower cover up) to show lower skin material (2024) in

contrast to upper cover and remaining structure (7075)


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10.6 Vertical Stabiliser and Rudder


Figure 39: C-27J Vertical Stabiliser FEM isometric view (left) and as per [5] (right)


Figure 40: C-27J Vertical Stabiliser shell elements (isometric view)


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Figure 42: C-27J Vertical stabiliser with partial skin removal showing internal elements (7075).


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10.7 Horizontal Stabiliser (starboard shown; port by similarity)


Figure 43: C-27J Horizontal Stabiliser FEM isometric view (upper) and as per [5] (lower)


Figure 44: C-27J Horizontal Stabiliser shell elements (isometric view)


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Figure 45: C-27J Horizontal Stabiliser beam elements (isometric view)


Figure 46: C-27J Horizontal stabiliser with partial skin removal showing internal elements (7075).


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10.8 Horizontal Stabiliser – Elevator


Figure 47: C-27J Elevator FEM isometric view (upper) and as per [5] (lower)


Figure 48: C-27J Elevator shell elements (isometric view)


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Figure 49: C-27J Elevator beam elements (isometric view)


Figure 50: C-27J Elevator with partial skin removal showing material allocation (7075).


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10.9 Cargo Door


Figure 51: C-27J Cargo Door FEM isometric view (upper) and as per [5] (lower)


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Figure 52 – C-27J Cargo Door shell elements (isometric view)


Figure 53: C-27J Cargo Door beam elements (isometric view)


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Figure 54: C-27J Cargo Door with partial skin removal showing material allocation. Outer skin:

2024 (blue), interior structure 7075(red), and attachment hooks 4130 (green)


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10.10 Cargo Ramp


Figure 55: C-27J Cargo Ramp FEM isometric view (upper) and as per [5] (lower)


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Figure 56: C-27J Cargo Ramp shell elements (isometric view)


Figure 57: C-27J Cargo Ramp beam elements (isometric view)


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Figure 58: C-27J Cargo Ramp with partial skin removal showing material allocation. Outer skin: 2024 (blue), interior structure 7075(red), and ramp hooks 4130(green)


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11. Conclusions and Recommendations for Future Work

This user manual has been created for the DST C-27J NASTRAN Finite Element Model designated as version C-27J_GFEM_DSTG-v3.0. The purpose of this manual is to provide sufficient detail to allow users to understand how the model is organised, to understand its limitations, and to understand how to best incorporate changes into the model. This manual is an aid to the use of the C-27J Global FEM, in particular to support on-going and future wing fatigue test (WFT) interpretation. Several areas have been identified as useful enhancements to the model and are recommended for future work in [6].


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12. References

1. Nastran FEM: C-27J_GFEM_DSTG-v3.0, ZIP File Format, 22 November 2017. Objective File ID: AV14738798

2. Opie, M., RELEASE OF THE DST GROUP C-27J FINITE ELEMENT INTERNAL LOADS MODEL VERSION 1.0, DST Minute, April 2016. Objective File ID: AV12755299

3. Opie, M., Review of the Alenia C-27J NASTRAN Finite Element Internal Loads Model, DSTO Minute, File B2/129/PT4,19 April 2013.Objective File ID: AV5654215

4. DGTA SOR, File Ref: U5110698, C-27J FY 13/14 DSTO Support to ASI-DGTA, Issue 1, 18 March 2014

5. C-27J Structural Repair Manual Objective File ID 14797650

6. Opie, M and Hadcroft, D., C-27J FEM Version 3.0 Enhancement and Verification, DST-Group-TN-1686, October 2017, Objective File ID: AV15170326

https://drms-edinburgh2/id:AV5654215/document/versions/published



UNCLASSIFIED

UNCLASSIFIED

DEFENCE SCIENCE AND TECHNOLOGY GROUP DOCUMENT CONTROL DATA

1. DLM/CAVEAT (OF DOCUMENT)

2. TITLE

C-27J NASTRAN Global Finite Element Model User Manual C-27J_GFEM_DSTG-v3.0

3. SECURITY CLASSIFICATION (FOR UNCLASSIFIED LIMITED RELEASE USE (U/L) NEXT TO DOCUMENT CLASSIFICATION)

Document (U) Title (U) Abstract (U)

4. AUTHOR(S)

Michael Opie and Damion Hadcroft

5. CORPORATE AUTHOR

Defence Science and Technology Organisation 506 Lorimer St Fishermans Bend Victoria 3207 Australia

6a. DST GROUP NUMBER

DST-Group- TN-1700

6b. AR NUMBER

AR-017-020

6c. TYPE OF REPORT

Technical Note

7. DOCUMENT DATE

November 2017

8. OBJECTIVE ID

9.TASK NUMBER

AIR 07/384

10.TASK SPONSOR

ASI-DGTA

11. MSTC

Airframe Technology and Safety

12. STC

Structural and Damage Mechanics 13. DOWNGRADING/DELIMITING INSTRUCTIONS

14. RELEASE AUTHORITY

Chief, Aerospace Division 15. SECONDARY RELEASE STATEMENT OF THIS DOCUMENT

Approved for public release OVERSEAS ENQUIRIES OUTSIDE STATED LIMITATIONS SHOULD BE REFERRED THROUGH DOCUMENT EXCHANGE, PO BOX 1500, EDINBURGH, SA 5111

16. DELIBERATE ANNOUNCEMENT

No Limitations 17. CITATION IN OTHER DOCUMENTS

Yes 18. RESEARCH LIBRARY THESAURUS

C 27J, aircraft, stress analysis, Finite Element Model, FEM, FEA, Verification, Nastran , Royal Australian Air Force, AIR8000 19. ABSTRACT

The Royal Australian Air Force (RAAF) has commenced a fleet acquisition of C-27J aircraft (AIR8000 Phase 2) to support RAAF tactical airlift capability requirements. As part of the Structural Substantiation Program, a global Finite Element Model (FEM) of the C-27J airframe was obtained from the Original Equipment Manufacturer Alenia Aeronautica. A global airframe FEM is an important supplementary tool in support of current and future RAAF C-27J structural integrity management. In the present user manual, detailed descriptions of the various sub-models constituent to the global model, including guidelines on their use are provided. Pending experimental validation, the enhanced and verified C-27J Global FEM is a linear elastic internal loads model that will be a useful tool in providing global loads results such as wing tip displacements, field stresses, running loads, connection forces, and other potential uses as described within the body of this document. The user manual provides important information on how to use and update the model, and the limitations associated with it.

C-27J NASTRAN Global Finite Element Model User Manual C-27J … · 2018-05-09 · UNCLASSIFIED. UNCLASSIFIED. C-27J NASTRAN Global Finite Element Model User Manual C-27J-DST-v3.0

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