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1 Initial Thesis Report for the Structural Optimisation of the Fuselage for the ADFA SAE Aero Design UAV

UNSW@ADFA, 2008

Structural Optimisation of the Fuselage for the ADFA SAE Aero Design UAV

Christopher John Kourloufas

ZACM4050 University of New South Wales

Australian Defence Force Academy School of Aerospace, Civil and Mechanical Engineering

Canberra, ACT 2600, Australia

The aim of this report is to present a review of the literature regarding the structural optimisation of aircraft fuselage structures as well as outline the direction of the Thesis project undertaken by the author. The Thesis project involves the optimisation of a fuselage structure for greatest specific torsional rigidity and specific flexural rigidity. The benchmark will be the remote-controlled UAV developed to fulfil the 2008 SAE Aero Design Competition specifications. Various alternative methods of fuselage construction will be investigated in order to derive the best solution.

Nomenclature σcr = critical stress for a thin sheet [MPa] K = non-dimensional coefficient that depends on the conditions of the edge restraint and shape of a plate E = Young’s modulus [GPa] t = thickness of a plate [m] b = width of a plate [m] σ = stress [Mpa] M = applied moment [Nm] I = moment of inertia [kgm2] x = perpendicular distance from the neutral axis [m] T = applied torsional load [Nm] A = cross sectional area [m2] q = shear flow [N/m]

!

" = angle of twist due to applied torsional load [degrees] L = distance to applied load from origin [m] J = polar moment of inertia [m4] G = shear modulus of elasticity [GPa] v = deflection of beam due to applied load [m] x = distance along beam [m] P = applied bending load [N] CAD = Computer Aided Design CATIA = Computer Aided Three-dimensional Interactive Application UAV = unmanned aerial vehicle Cobra = UNSW@ADFA entry to the 2008 SAE Aero West Competition SAE = Society of Automotive Engineers


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I. Introduction The function of the fuselage of an aircraft is to transmit and resist the applied loads; provide an aerodynamic

shape and to protect the payload and systems of the aircraft. The fuselage supports major concentrated loads such as wing and tailplane reactions, undercarriage reactions and inertia forces (Megson, 1999, pp.223-225). It is the role of the designer to manage these loads whilst considering the impact of the design on aspects such as weight. Weight optimization of the fuselage structure is an important phase of the design process as it impacts the performance of the aircraft. By optimizing the specific rigidity of the fuselage, weight can be reduced and thus payload or fuel volume may increase (Niu, 1999, p.5). Other drivers for optimizing the structural rigidity of the fuselage can include increased maneuverability and reduced acquisition, fabrication and assembly costs (Schetz, 2004, p.3).

Structural optimization has great relevance in model UAV design due to the model constraints of size, weight and power. The aim of the Thesis project, which this initial report describes, is to optimise the fuselage structure of a model UAV for greatest specific torsional rigidity and specific flexural rigidity. The benchmark will be the remote-controlled UAV developed, by this student and three other students, to fulfill the 2008 SAE Aero Design Competition specifications. This project will investigate alternate designs and materials to achieve this aim. The results of the project will be verified through experimental and analytical tests. A full statement of the client brief, which highlights the project outcomes, is attached in appendix A.

This Initial Thesis Report will include a literary review of the modern methods of fuselage design, an outline of the design methodology to be used in this project, the work completed to date, and finally, a description of the management scheme and methods to be used to complete the project.

II. Review of fuselage design literature

A. Fuselage design of modern aircraft The convention for modern aircraft fuselage design is to utilize a semi-monocoque structure. This structure is

comprised of a cylindrically shaped skin shell, which is stiffened by elements such as frames, bulkheads, stringers and longerons (Megson, 1999, p.225). The longitudinal stiffening elements are stringers and longerons and the transverse stiffening elements are frames and bulkheads. The longerons carry the majority of the axial loads resulting from the bending moments, and the skin carries the shear from the applied transverse and torsional forces (Niu, 1999, 376). This skin-stringer panel construction is a very efficient structure for aircraft as is provides a continuous surface at a very low weight (Niu, 1999, 141).

Applied loads are transmitted via shear forces, through the connections between the structural elements. These connections can be fasteners, adhesive bonds or welds. The applied forces are transmitted usually along the length of such connections to the reinforcing member. The applied force is limited through the resistance to bending or twisting of the reinforcing member (Niu, 1999, 141).

Numerically, for a thin sheet, the critical compressive stress is given by,

!

"cr

= KEt

b

#

$ % &

' (

2

(1)

Thus, as (b/t) increases, the critical stress decreases. The methods for counteracting this effect are to increase sheet thickness or to use reinforcing stiffeners. Stiffeners used are usually closely spaced and approximately the same thickness as the skin. This enables the structure to be analysed as one unit (Megson, 1999, p.176).

B. Failure in semi-monocoque structures The semi-monocoque structure will become unstable in one of three ways: skin instability; panel instability or

general instability. Skin instability is related to the stress relation given in Eq. 1 whereby a thin sheet will buckle under applied stresses. This is alleviated by the total effect of the longitudinal reinforcing members of the structure such as longerons and stringers. Fuselage rings and frames divide the stringer longerons and their corresponding skin into lengths called panels. Panel instability arises due to the longitudinal reinforcement acting as columns with an effective length equal to the frame spacing. This failure is alleviated by ensuring the frames and bulkheads are sufficiently stiff to support the longitudinal reinforcement in column action (Niu, 1999, 378). Finally, general instability occurs over two or more panels and is not confined between two adjacent frames. This type of failure is


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mitigated by ensuring bulkheads and frames are spaced at the nodes of bending and are sufficiently rigid (Hoff, 1986, p.105).

C. Analytical modeling of fuselage structures Many methods have been developed for the analysis of the semi-monocoque structure of an aircraft. Methods

range from “back-of-envelope” calculations using basic mathematical equations found in elementary texts; to the simultaneous solving of thousands of differential equations by computer. Research has been conducted in showing the difference in precision of analytical methods that vary in complexity (Boni, 2004). This research has found that a high level of accuracy is achieved through finite element analysis. Finite element analysis allows for a greater model development and accounts for realistic load cases and structural interactions. This translates to greater depth of investigation of the structural behaviour (Boni, 2004). Finite element analysis is conducted with the use of computational software such as ANSYS or CATIA (ANSYS, 30 April 2008) (Dassault Systemes, 29 April 2008).

D. Fuselage shape The shape of the fuselage has an impact on the aerodynamics of the overall aircraft (Raymer, 2006, p.483). The ideal fuselage shape would contribute to lift and exhibit minimum drag characteristics whilst having a stablilising effect on the overall aircraft. This is a difficult problem to solve, as a flow analysis, via experiment or computational means, must be undertaken. Analytical fuselage optimisiation has been investigated by many researchers and involves a high level of analysis (Peigin, 2006). An aerodynamic analysis of the fuselage is beyond the scope of the Thesis project, however, the effects of fuselage shape should be considered. To satisfy this need of an aerodynamic shape, research has been conducted in the area of low-drag bodies. A simple low-drag shape has been presented by Simons in the text, Model Aircraft Aerodynamics, which would serve as an ideal starting point in the conceptual design of a model fuselage.

III. The Cobra model UAV The Cobra model UAV was design to the requirements set out by the 2008 SAE Aero Design Competition rules

(SAE, 23 April 2008). Four students have developed the Cobra since November 2007 and have incorporated the Cobra into their final year Thesis projects. The project described in this initial report aims to study the structural optimization of model UAV fuselage design and apply the findings to the Cobra model UAV.

A. Customer requirements Through analysis of the 2008 SAE Aero West rules (SAE, 23 April 2008) a list of requirements was produced

that dictated the physical dimensions of the aircraft as well as performance. Such requirements include the need for the combined length, width and height to remain less than 175 inches and the need to include a cargo payload area 5x5x10 inches in size. A detailed list of requirements is found in appendix B. To achieve these requirements, the design methodology adopted for the Cobra was to optimize aerodynamic efficiency and pursue highest lift whilst remaining within the specified physical dimensions. The implications of this methodology on the fuselage design were to minimize weight while retaining torsional and flexural rigidity, length restrictions and a need for a low drag profile.

B. Initial Concept The initial concept selected was a canard aircraft, as this type of aircraft would easily satisfy several

requirements. Firstly, the canard produces lift that adds to the lift produced by the wing (Nelson, 1997, p.52). Secondly, the nature of the canard design maximizes usable fuselage volume. This derives from the substitution of a canard instead of a long empennage for longitudinal stability. This also allowed for a shorter fuselage. Another favorable aspect of the canard design is the position of the centre of gravity – between the two lifting surfaces (Anderson, 2004, p.531). This allowed the payload to be positioned in a location that would have minimal impact on the stability and control of the aircraft. This also reduced the need for ballast to enhance longitudinal stability which reduces aircraft empty weight. The concept is depicted in Figure 1


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Figure 1 - Initial Cobra concept

C. Preliminary Design A novel method for sizing a canard model aircraft was developed which took two different approaches to arrive

at the final aircraft size. A safety factor of 11.81 inches (300mm) was included in the overall dimensions in each method of sizing. The first approach was to use the historical data (Jane’s, 23 April 2008) of current canard aircraft in operation. This historical data is found in appendix C. The dimensional data of six canard aircraft were used to plot trends between length versus wingspan, wing area versus canard area and wingspan versus canard span. It was found that all three comparisons displayed strong trends and would serve as a basis for a first estimation of the model dimensions. This method worked by setting the height off the ground and simultaneously solving the length and wingspan using the equations:

total dimensions = length + wingspan + height + safety factor (2)

wingspan = 1.34*length. (3)

The iteration that was selected was one with a reasonable height taking into account propeller clearance. The fuselage length was found to be 62.20 inches (1580mm)

The second approach was to set the height and length of the model by the real-world constraints and thus determine the wingspan. Such real-world constraints include propeller clearance, propeller diameter and systems lengths. Fuselage length was estimated to be 43.70 inches (1110mm), height from the ground was estimated to be 16.93 inches (439mm) and wingspan was estimated to be 102.56 inches (2600mm).

Finally, a third set of model UAV dimensions were developed as a compromise between the two previous methods. At this stage, 17.72 inches (450mm) was set as the optimum model height and an intermediate model length of 55.11 inches (1400mm) was selected. These constraints allowed the wingspan to be 90.55 inches. The detailed data for these iterations, and the final aircraft dimensions are found in appendix D.

D. Detailed Design The design of the superstructure was comprised of a keelson beam running the length of the fuselage, bulkheads

at critical points along the fuselage and longerons joined to the bulkheads. The fuselage was designed to gain most of its bending strength from the composite keelson beam that runs along its length. Plywood bulkheads spaced at critical positions gave the fuselage cross-sectional structure and wooden longerons gave torsional stiffness. The fuselage assembly was then covered in plastic shrink-film, thin balsa sheeting or a combination of the two to give extra torsional stiffness.

The design load factor was set a three times the force of gravity for all structural calculations as this is a reasonable estimation of the loads experienced in high stress conditions such as landing. During the design of the Cobra, basic static load cases were analysed at this load factor to determine the size of the superstructure. The two conditions that governed the size of the superstructure were bending during flight and torsion loads due to landing forces. It was determined through manual calculations, as well as computer modelling, that Tasmanian Oak alone would suffice for torsion and bending stresses. Three carbon fibre/epoxy resin layers were added to the keelson beam as an extra safety factor. This would increase the stiffness of the beam and would help resist bending and torsional loads.

E. Construction of the Cobra fuselage Due to the time constraints placed by the SAE Aero Design competition, a detailed structural analysis was not

conducted. It was found during the manufacture process that the torsional deflection caused by the wings was


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extremely high. To reduce this deflection, the entire fuselage skin was made from balsa sheet, which increased the critical stress value, as shown in Eq. 1. As well as this, all joins and fittings were made to a high tolerance so that there was minimal freedom of movement of the joins. The combination of these improvements reduced the torsional deflection significantly.

The materials used in the construction of the fuselage have been described below in Table 1.

Material Component(s) Balsa wood Longerons, skin, wing box Aluminium Main landing gear strut, canard box, wing box Plywood Bulkheads Oak Keelson beam Bi-directional carbon fiber/epoxy resin composite Keelson beam, wing box Heat Shrink plastic covering film Skin Plastic resin Engine mount Spring Steel Nose landing gear

Table 1 - Materials used in the Cobra fuselage

The materials used were easily accessible and simple to work with. The Cobra, pictured in Figure 2, was constructed with the use of hand tools, and simple machinery such as a band saw, lathe and drill press. A construction methodology and schedule was developed during the design phases. However, the complexity of the construction was underestimated, and therefore the schedule and methodology were inaccurate. This flaw was primarily caused by a lack of experience in model construction. To rectify this, the planning of the construction was conducted at the sub-component level just before the component was to be built. This enabled greater visualisation of the tasks involved and reduced the impact of inexperience. The lessons learned during this phase were recorded so that they may be applied to the construction of future model aircraft.

IV. Thesis Project – Structural Optimization of the Model UAV Fuselage

A. Customer requirements A Quality Functional Deployment (QFD) was conducted for the thesis project based on the client brief and the

2008 SAE Aero Design Competition Rules. The QFD identified the essential and desired customer requirements as found in Table 2. The detailed QFD is attached in appendix E.

Essential Requirements Desired Requirements (in order of importance)

A study of the optimization of a UAV fuselage The ability to be tested experimentally for rigidity Fuselage designs that implement the results of the study Technical drawings of the fuselage

A construction methodology Results to accompany the study from analytical and experimental tests Accessibility for tooling and construction A completed Thesis report on 20 Oct 2008 Internal volume to install systems A number of constructed fuselages for testing Integration with wing and canard construction

An engine mount A payload capability of dimensions 5 x 5 x 10 inches The capability to inspect all systems An undercarriage construction Vibration analysis

An aerodynamic shape

Table 2 - Thesis Project Essential and Desirable Requirements

Figure 2 – Cobra fuselage construction


UNSW@ADFA, 2008

The hierarchy of desired requirements gives clear direction for the Thesis project and thus time can be allocated relative to the importance of a requirement.

B. Critical discussion of modern fuselage design methods and application to model UAV design As discussed earlier, the semi-mocoque design for an aircraft fuselage is the modern convention. This method

can be employed in the design of a model UAV with some limitations. Firstly, most materials used in the manufacture of a model UAV would not have the same relative proportions that full-sized aircraft material possess. For example, the skin thickness relative to the overall dimensions of a full-sized aircraft as compared to a model UAV would differ by a considerable amount. That is, the skin of a model UAV would be much thicker than it needed to handle the loads experienced by the structure.

This leads to the second point of discussion, producability. It would be a laborious and tedious task to fabricate a scaled-down version of a monocoque fuselage for this model UAV. The risk of damage to components during manufacture would increase due to the fragility of the scaled-down structure. Thus, the modern convention for fuselage design must be adapted to account for the scaling issues that arise from model aircraft construction.

V. Model UAV fuselage design methodology

A. Previous design input As the configuration of the model UAV has already been determined throughout the design phases of the Cobra

UAV, there is knowledge of the position of features such as the wings, canard and the undercarriage. With this information and the customer requirements, a methodology can be established.

B. Design considerations A critical design consideration for the model UAV fuselage will include the reinforcing of sections where

appropriate measuring equipment and experimental jigs can be fixed to in order to measure the torsional and flexural rigidity. Without this capability, there will be no method for experimentally verifying the design of the fuselage. Also, integration with the wings and canards will be of high importance, as the loads from these lifting surfaces must be transmitted throughout the fuselage in order to reduce stress within the structure.

C. Model fuselage conceptual design process The conceptual design methodology of the model UAV fuselage will follow a modified version of the design

general requirements presented by M.C.Y. Niu. The model UAV conceptual design process is as follows (Niu, 1999, p. 379):

1) Chose an aerodynamically smooth shape (outline)

2) Provide space for and position internal systems and cargo

3) Position hardpoints for attachment of experimental jigs

4) Make arrangements for undercarriage placement

5) Position bulkheads

6) Position frames

7) Consider accessibility for tooling and maintainability – cutouts in fuselage

8) Reinforce cutout regions

By following this methodology, a fuselage concept can be produced and the internal structure can be developed.

The design of the internal structure will concentrate on dispersing the applied aerodynamic and ground loads throughout the fuselage. This will be achieved through the use of a stringer-skin arrangement and bulkheads.

D. Preliminary design of the model UAV fuselage Preliminary design will consist of analysing the loads through basic numerical calculations of a simplified

fuselage structure. According to M.C.Y. Niu, the design of fuselage structures involves two steps: solving the stress


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distribution under all load conditions and checking the structure can sustain these stresses (Niu,1999,p.377). Therefore, the load experienced by the model UAV conditions must first be defined (Table 3).

Bending

(about the centre of gravity) Torsion

(about the fuselage centroid) On the ground with maximum payload – moments due to the engine, payload and reactionary forces

One wheel landing (MLG) – torsion generated between MLG and wing box

Maximum lift generated by wing and canard - moments experienced between wing and canard

Maximum cross roll – moments due to wing and canard rolling in opposite directions

Table 3 - Load conditions for initial stress determination

To check that the structure will be able to sustain the stresses generated by the load conditions listed in Table 3, some basic mechanics of materials theory will be applied. The parameters checked and formulas used are listed in Table 4 (Beer, 2005).

Parameter Formula applied

Skin thickness (Eq. 1)

Frame/bulkhead spacing

FrameEIMDL

)(16000

2

= (Eq. 4)

Bending in lateral plane

!

"z

=M

y

Iyy

x (Eq. 5)

Torsion about the longitudinal axis

!

T = 2Aq (Eq. 6)

Table 4 - Parameters checked and formulas used

Once the preliminary design has been locked in, construction of the first test fuselage will take place. While the first test fuselage is being constructed, an analytical model of the preliminary design will be developed as part of the detailed design phase. This is because detailed design will require a high level of analysis and to conduct this analysis by hand will be a task too complex for the scope of this Thesis project. A discussion of the analytical model methodology is included in the following sections. Finally, the detailed design process will involve an investigation on the properties of various alternative materials to be utilised in the construction of the model UAV fuselage. Such material will include paper/cardboard, plastics and fiberglass.

E. Experimental analysis The application of the theory outlined will be validated through

experimental analysis of the fuselage under the stress conditions described in Table 3. This will be achieved by clamping the fuselage at the appropriate end and loading the other to measure the deflection of the structure. Effort will be made to decouple the bending and twisting motions so that pure bending and pure torsion are closer to being achieved. These tests will be conduced using existing laboratory equipment, for example the experiment apparatus pictured in Figure 3. Apparatus such as this will be modified in order to accommodate the fuselage being tested. Both the Cobra and the new model UAV fuselages will be tested and results compared. The analytical model will also be compared to the experimental data to determine whether the problem was modeled correctly.

The theory used to analyse the fuselage construction will involve the twisting theory and the cantilever beam bending theory as detailed below.

2

!"

#$%

&=

btKEcr'

Figure 3 – Experimental apparatus


UNSW@ADFA, 2008

Twisting: (Eq. 7)

!

Bending:

!

EId

2v

dx2

= "Px (Eq. 8)

The bending equation (Eq. 8) will be double integrated, with the appropriate boundary conditions, to determine

the displacement v of the beam. The experiment will involve incrementally increasing the applied load and measuring the deflection. A deflection versus applied load graph will be plotted and the slope of the graph will give the mechanical properties of the fuselage as per the relations in Eq. 7 and Eq. 8.

F. Analytical modeling The CAD package, CATIA will be utilised to model and analyse the fuselage design. The reason for this is the

capability of the software to be used for the design of composite parts and the finite element analysis of those parts (Dassault Systemes, 29 April 2008). As only static cases will be analysed, the level of detail that the CATIA software is capable of will suffice. The software allows for control of the mesh size, boundary conditions and applied loads. Meshing is the process of representing a physical domain with a number of finite elements. The mesh can be manipulated to deliver more accurate results at the cost of processing time due to complexity (Hutton, 2003,p.4).

The use of a single software package for pre and post-processing is favourable as there will be no compatibility issues during the analysis. That is, the model will firstly be developed within a design workbench of the program, and then transferred to an analysis workbench, within the CATIA program, to conduct the finite element analysis. Finally, the bidirectional associativity that exists between the various workbenches of the CATIA software allows for the carry-through of design changes between those workbenches (Tickoo, 2006, p. xx). This feature will allow efficient use of the time allocated for analytical modeling.

G. Project management An initial project plan was drafted and submitted to the customer, which included a project task outline, task

breakdown structure, a project schedule and a milestone chart. The milestone chart is included in Table 5, which describes the date of delivery of major Thesis project components.

Milestone Target Date

1. Submission of Project Plan documents. 04 Apr 08

2. Submit Initial Thesis Report 30 Apr 08

3. Construct Initial Fuselage 13 Jun 08

4. Test Analytical Model 22 Aug 08

5. Test Final Fuselage Model 19 Sep 08

6. Submit Thesis Report 20 Oct 08

Table 5 - Project milestones

The project will be run in accordance with the Gantt chart produced during the planning phase. The project management documentation has been attached in appendix F

VI. Summary In conclusion, this Initial Thesis Report has defined the necessity for structural optimization of the aircraft

fuselage, previous methods of such optimization and the application of these methods to this Thesis project. This literary review forms the context with which the Thesis project will be based upon and gain its direction. The report has also outlined the work conducted by the author on this topic as well as the methodology that will be used to conduct the remainder of this project. The Thesis project will require design, manufacture, analytical optimization and experimental analysis in order to optimise the model UAV fuselage structure for greatest torsional and flexural specific rigidity.

!

" =L

JGT


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