Rear Fuselage structure

7/28/2019 Rear Fuselage structure

http://slidepdf.com/reader/full/rear-fuselage-structure 1/17

DESIGN OF A REAR FUSELAGE

Figure A.1 shows the elevation of Sirius Aircraft. It is required to carry out the detailed structural design

of the portion of the rear fuselage between the sections AA and BB.

Figure A.1 Sirius

A.1 Specification

The required flight envelope for this particular aircraft is shown in Figure A.2 where



Note also that airworthiness requirements specify that since the point D2 lies on the axis.

Further requirements are that:

(i) at any point in the flight envelope an additional pitching acceleration is given by

(ii) for asymmetric flight an angle of yaw given by

A.2 Data

Preliminary design work has produced the following data.

AIRCRAFT

Fully loaded weight = 7.56 k.N.

Moment of inertia of fully loaded aircraft about the center of gravity = -0.0155.

The engine has a maximum horsepower of 115hp and the propeller efficiency is 85%.

WING

The wing has a span of 9.75m and gross area of 11.89m2. Its aerodynamic mean chord, and

the variations of the moment and lift coefficients with incidence are shown in Figure A.3.

Figure A.3 Wing characteristics



Also, the pitching moment coefficient is given by

TAILPLANE

The tailplane has a span of 2.65m and a gross area of 1.46m2; the position of the aerodynamic center, P,

of the tailplane is shown in Figure A.1 which is 20% of the tail’s mean aerodynamic chord.

Due to the asymmetry of the slipstream induced by the yaw asymmetric loads are induced on the

tailplane. These loads, upwards on one side and downwards on the other, result in a torque given by

√

FIN

The fin has a height of 1.143m, an area of 1.04m

2

and an aspect ratio of 1.3. Also, it may be shown thatthe lift-curve slope, , of the fin is given by

The position of the centre of pressure of the fin depends upon the geometry of the pressure

distribution. Calculations show that the centre of pressure is 0.45m above the axis of the rear fuselage

and a distance of 2.78m aft of the section AA.

A.3 Initial Calculations

FLIGHT ENVELOPE

The positive stall curve in the flight envelope of Figure A.2 is found from the basic aerodynamic wing

theory is given by

Then

⁄ ⁄

FIN LIFT-CURVE SLOPE



A.4 Balancing out Calculations

The tailplane and fin loads corresponding to the various critical points in the flight envelope will now be

calculated so that, subsequently, values of shear force, bending moment and torque acting on the rear

fuselage may be determined.

Also

A first approximation for the wing lift, L, is obtained by neglecting the tailplane lift, P,

The angle of yaw is,

The coefficient of lift and total pitching moment coefficient is,

so that

The additional pitching moment acceleration is,

The tail load

.

The torque produced by the asymmetric loading on the tailplane is given by,

√



The load on the fin caused by the yawed flight is given by,

The torque produced on the fuselage by this fin load is 663.3652 × 0.45 = 298.5143 Nm. The total torque

on the rear fuselage is therefore given by

A.5 Fuselage loads

The dimensions of the portion of the rear fuselage to be designed are given in Figure A.1.

FUSELAGE SECTION

For the Sirius aircraft the fuselage is unpressurized so that the frames will not support significant loads.

However they will be required to maintain the fuselage shape but may therefore be nominal in size,

suitable frame sections will be suggested later. The combination of stringers and skin will resist the

shear forces, bending moments and torques produced by self-weight and aerodynamic loads. For this

purpose a circular cross-section will meet the design requirements of the aircraft and be simple to

fabricate and design.

Figure A.4 Stringer arrangement in rear fuselage.

Figure A.4 shows a possible section. Twenty-four stringers arranged symmetrically, each having

the same cross-sectional area.



MATERIAL

An aluminium alloy will be used for both stringers and skin and has the following properties:

SELF-WEIGHT

In a conventional single-engined aircraft of the type shown in Figure A.1 it is usual to assume that the

fuselage weight is from 4.8% to 8.0% of the total weight and that the weight of the tailplane/fin

assembly is from 1.2% to 2.5% of the total weight. It will be further assumed in this case that half of the

fuselage weight is aft of the section AA and that the weight distribution varies directly as the skin

surface area.

Figure A.5 Rear fuselage sections.

For ease of calculation the rear fuselage is assumed to taper uniformly as shown in Figure A.5; CC is a

section midway between AA and BB. The total skin area is given by



Also the center of gravity of the tailplane/fin assembly has been estimated to be 1.64m from the sectionAA on a line parallel to the fuselage centre line.

SHEAR FORCES AND BENDING MOMENTS DUE TO SELF-WEIGHT

At the section AA

*

+

At the section CC

[ ]

*

+

At the section BB

[ ]

*

+

TOTAL SHEAR FORCES, BENDING MOMENTS AND TORQUES

The values of shear force, bending moment and torque at the sections AA, BB and CC will now be

calculated.

Section AA

The shear force due to the self-weight and tail load is,

()



The shear force due to the fin load is,

The bending moment due to the self-weight and tail load is,

The bending moment due to the fin load is,

()

The torque due to asymmetric flight and the fin load is,

Section CC

()

()

Section BB

()

()

A.6 Fuselage design calculations

In this case, Elastic design will be used. The working, or allowable, stresses are then

The proposed fuselage section is circular as previously shown in Figure A.4. The design process is

required to produce suitable stringer sections and a skin thickness. Suppose that each stringer (or boom)has a cross-sectional area B mm

2and that the skin thickness is t mm. The idealized fuselage section is

shown in Figure A.6.



Figure A.6 Idealized fuselage cross-section

STRINGER SECTIONS

The direct stress in each stringer produced by bending momentsM x and My is given by,

where

Section AA (diameter

)

At stringer 7,

Then

Similarly



Section CC (diameter)

Section BB (diameter)



From the above it can be seen that the maximum direct stress at each fuselage section occurs in

stringer 10. Also the stress in stringer 10 (and all other stringers) is lower at section CC than at section

AA and lower at section BB than at section CC. Therefore if fuselage frames are positioned at each of

these sections lighter stringers may be used between CC and BB than between AA and CC. An additional

frame will be positioned midway between AA and CC and between CC and BB, and will be slotted to

allow the stringers to pass through. The arrangement is shown diagrammatically in Figure A.7.

The allowable direct stress in a stringer is . The maximum direct stress in stringer 10

at the section AA is . The required stringer area of cross-section is then given by

Figure A.7 Arrangement of fuselage frames



SKIN THICKNESS

The fuselage cross-section is subjected to shear loads and along two perpendicular axes of

symmetry.

Since, is constant round the fuselage section and ,

Consider the action of only. Then

Referring now to Figure A.6

From symmetry so that

giving



Then

Similarly

Consider now the action of only.

From symmetry

The shear flow produced by the applied torque is given by

The total shear flow in each of the panels 76 to 21 is then



Section AA (diameter)

Then,





The skin will be of constant thickness so that the maximum shear stress in the skin will occur in thepanel in which the shear flow is a maximum. This, from the above, is 1.06 N/mm in panel 43 at section

BB.

RIVET SIZE

Skin/stringer rivets

Section AA (diameter)







From the above it can be seen that the maximum load on the rivets occurs at section BB in

stringer 9 and is 0.28 N/mm. The number of rivets/m is given by

Documents

Rear Fuselage structure