[nasa] - Nasa-cr-197172 Cabin-fulsage-wing Structural Design Concept With Engine Installation

NASw-4435

N_s.a,_C_,°l-91112

Cabin-

Fuselage-Wing Structural Design with Engine Installation

Concept

SOW #421F93ADP01-7 6 December 1993

AE 421/02/A Lead

Engineer: Scott Ariotti Team Members:

M. Garner, A. Cepeda J. Vieira, D. Bolton Submitted to: Dr. J. G. Ladesie

(NASA-CR-197172) CA3IN-FUSELAGE-WING DESIGN CONCEPT INSTALLATION

WITH (USRA)

N95-12993

STRUCTURAL ENGINE 87

p

Unclas

G3/05

0026160

Table

of Contents

1.0

Project Summary ........................................................

Page

1

2.0

Description

Page

1

3.0

Loads

4.0

Structmal

5.0

Manufacturing

6.0

Weight

7.0

Conclusions

of the Design

and Loadings

..................................................

......................................................

Substantiation

..................................................

and Maintenance

Estimation

Provisions

....................................

......................................................

.............................................................

Page 4 Page

11

Page

15

Page

17

Page

19

List of Figures and Tables

Figure 1

Wing Lift Distribution

Figure 2

Wing Shear Diagram

Figure 3

Wing Moment Diagram

Figure 4

Front Spar Shear Diagram

Figure 5

Front Spar Moment Diagram

Figure 6

Rear Spar Shear Diagram

Figure 7

Rear Spar Moment Diagram

Figure 8

Wing Torsion Diagram

Figure 9

Front Spar Section Modulus

Figure 10

Rear Spar Section Modulus

F93-2A-102-7

Fuselage Strucuaal Layout and Details

F93-2A-103-7

Fuselage Skin Panel Layout

F93-2A-104-7

Wing Structtwal Layout and Dotail,¢

F93-2A-105-7

Control Systems Layout

F93-2A-106-7

Wing Skin Panel Layout

F93-2A-107-7

Wing Access Panel Layout

F93-2A-1R3

1. Project

system

Summary

The purpose

of this project

is to provide

a

routing.

fuselage

The difficulties

design

include

expanding

fuselagestructural assembly and wing structural

cabin to accept a third occupant

design thatwillbe ableto withstandthe given

configuration

operationalparametersand loads provided by

safety.

Federal AviationRegulationPart23 (FAR 23) and

will move

the Statement of Work (SOW).

aft to compensate

The goal isto

and providing

By adding

transfer the appliedloadsthrough the most efficient

structure

load path. Areas of producability and

The diffculfies

maintainability of the structure willalsobe

design

addressed. All of the structural members willalso

capacity

meet or exceed the desiredloading criteria, along

torsion

with providingadequatestiffness, reliability, and

great

fatiguelifeas statedin the SOW.

problem

will result

need to be made for controlsystem routingand

Finally,

an adequate

cabin heating/ventilation. The goal of the wing

wing

structure and carrythrough structure isalsoto

effectively

provide a simple,lightweightstructure thatwill

2. Description

transferthe aerodynamic forcesproduced by the

2.1 Fuselage

Considerations

ample

the provisions

encountered resizing

and weight, produced

the CG of aircraft needs to be move

in the wing

the wing

inside

in relocating carrythrough

interface

will

loads

through

mount.

structural

for the increased

compensating

of stiffeners

transmit

in the moment.

of the engine

for a large

by the tail boom

and fuselage

for their

of a ring frame

for the new position

by placing the boom,

a this

the fuel tank. structure

for the

be designed

to

the fuselage.

Structure

The cabin fuselage

wing, tailboom,and landinggear.These forceswill

volume

for the difference

This required

number

in a staggered

so the engine

provide a durablelightweightstructure thatwill

in the

the fuselage

a third person

forward

include

encountered

be channeled through variousinternalstructures

stiff structure

sized forthe pre-determinedloadingcriteria. Other

under

considerations were to includespace for flaps,

FAR 23.

ailerons, fueltanks,and electrical and control

the use of various members

structure

that will maintain

the applied

provides

the proper

limit loads as indicated

The applied

loads

are distributed such

a

shape in the through

as longerons,

ring

F93-2A-11L3

frames,

bulkheads,

for bending longerons

moments,

Z - channels

formed

Lightening

and shear

assembly

C - channels

weight.

along

In addition,

the bottom,

to account

for torsion

The

emergency

stiffeners

produced

landing

removed

adequate

stiffness

to the nose

is installed

A series

this ring frame

out

while

2.4 Win_

directly

of longerons

to the aft bulkhead

for the engine during

to

The minimum

by determining

airfoil

that would

ingress

occupant

thickness

and egress.

for by using

occurs

approximately

of extruded

the Jungle

safety,

T-sections.

and

attached quadruple

airfoil

by NAS

determining

the maximum

chord

length,

of the chord.

the position,

several

for construction

from a series

to brake-formed

C - channels.

to use an aluminum

thickness 344 series

This built-up to the carrythrough shear, 4340

the given

of 25% to

the spar at 25%

I - beams

within

spar

the height to be a

it was decided

bottom

Aviation

allow

were considered

requirements

the front

the position

70% of the total

After

all loading

These areas are

of designing

over a range

placing

ideas

seating,

for human

aircraft control.

For the NLF 0414

web of increasing volumetric

need to assure adequate

maintaining

limits

extent

Spar and Lug

provide

Requirements

are

and the geometric

acceptable

began

leading

The seat adjustments

to

Structure

Front

thereby

seat belt connections

physique,

provide

comfort

Finally,

2.3 Spatial

provided

of travel

maximum.

gear

conditions.

and cabin

to pilot

under

erashworthiness

structure.

The process

is fabricated

seat, that is certified

and four point

the fuselage

2.4.1

The structure can be easily

The powerplant

the aft ring frame.

the Z - channels,

and is installed

to allow for easy access

assembly.

been

Assembly

composite

screws.

have

(JAARS)

current FAR 23 dynamic

conditions.

The nose cone assembly

flathead

Services

relative

by the occupants

Mount

of a fiberglass/epoxy

Radio

conditions,

are extruded

the sides.

in between

2.2 Nose Cone/Engine

from

flow.

under the floor and 2024-T3

used along

using

paths

holes are cut in the extruded longerons

to reduce

during

These provide

torsion,

of the fuselage

7075-T6 brake

and skin.

shear

capped

on top and

extruded

aluminum

member using

would

then be

machined,

steel lugs and shear bolts.

F93-2A-1R3 Usingthetotallift distribution, shearand

were almost identical,

thereby

bendingmoment diagrams weregenerated for the

of the spar caps to depend

entire wing.

and bending

V-n diagram

From the maneuvering 88% of these

the front spar as a result analysis.

From these

section

modulus

T-section section 344.

diagrams

was determined

Rear

based

and available

until

similarly

sizes of stock NAS the increased

gear a large

of inertia

section

modulus

section

modulus.

section

modulus

was chosen and varied member using

web type,

incorporating to a aluminum approximately

as a built-up

shear

lifting

load of the entire wing at the maneuvering

Since the wing was chosen to a NLF

0414

airfoil, the heightof the frontsparand rearspar

344-2

at the tip.

quadruple

T - section

This built up

to the carrythrough

shear,

4340

steel lugs

Structure structure was initially of the front and rear

of the engine

significantly

conditions.

The front

incorporated

of the hoop,

the carrythrough

riveted

of the aircraft.

was designed

connected

was

spar carrythrough

to the ftrewall

C - section

of the tail boom,

to accommodate the loading

two C - sections

spar carrythrough

(A) on the V-n diagram.

the required

extruded

spars, but due to the configuration

attached

the available

versus

344-32

as a continuation

modified

to take up 12% of the total

a NAS

The carrythrough

The rear spar

and was designed

the required

In the case of the available

Carrythrough

shear

airfoil

to provide

bolts.

and the location

chord on the

shear

for each section.

then be attached

spar caps

at the 67%

diagrams

(S) is plotted

to a NAS

T - sections web.

on the respective

and subsequentially

machined,

designed

For strength/weight

the design

as the spar cap at the root of the wing

would

of the rear spar was performed

the spar was designed

point

moment

2.4.3

spar.

moment

The spar caps needed

and shear

Lm,

to the front

was located

loads

80.

efficiency

attached

along

on the required

and landing

station

Spar and

The design

a required

The sizes of the spar cap

by the tail boom

AS was retained

equilibrium

at intervals

In order to accommodate

imposed

2.4.2

of a chordwise

were then chosen, modulus

(A) on the

loads were applied to

loading

the span and plotted.

point

allowing

to another

this change

together

and

The rear

with a half-hoop C - section

was necessary

on top

due to the

F93-2A-1R3

location

of the engine.

3. Loads

(Ref.

and Loadin2

3.1 Fuselage

the structure to comply

withstand

Bending,

The highest

bending

condition.

on

In order

with FAR Part 23, the structure must be

Both the bending

passengers,

the .IAARS

loads

through

to the wing main

calculation

is shown

spar. This

below.

M_,

= 727.2(10 _) Ib-in and the resulting

of safety

Engine

assumed

is less than the given

the engine

rivet

ultimate

stress of 4g ksi, and this results in a margin safety 3.2

of 0.363.

Safety

Harness

of

the

lbs

yields

similar

results

produced

mounts

by the engine

in lbs, and is reacted

to the ring frame.

the torque

spacing

into shear

must be determined

under

using the method 139).

through

This, in ram,

The skin thickness and

skin panels.

loads imposed

was

flow in the

to withstand

by the torque.

frame area of 1,773 in2, the resulting psi is well

with a

of 1.17.

to be 12,000

buckling stress

under

Torque

surrounding

This bending

was sized

inches. A similarevaluationof the rearoccupant's

translates

F_=Wa¢o, *39=1987 (3) FsG=59611bs MNG=Fz_G*d_=5961 (122 ) MSG=727 200 inlbs

this stress an

The resulting thicknessof the T - sectionis0.220

The torque

f_ = 29.3 ksi.

a bearing stress

D_t = 0.5in.

3.3

stress is

to FAR

conditions:

harness attachment

from the nose gear load are

must

f,a-- = 99 ksi

margin

moments

T - section

V = 10.91(103)

Fe= [3 *Wg÷3 *W_s] *3g =[3 (200)+3(30)].3 =207 O/bs Mc=f c*dc=207 0 (40) =82800inlbs

The bending

aluminum

following

of the

the floor structure

bolts have

In order to accommodate

extruded

system, and the nose gear

are reacted

and the longerons

moment

harness

of 18g's according

The attachment

of 99 ksi.

a 3g impact under normal landing

conditions.

moment

loads are imposed

shoulder

an impact

Part 23.

during the landing

able to withstand

landing

The JAARS

Dwg. F93-2A-104-7)

For a ring stress

of 136

the F_t of 1778 psi prescribed outlined

the

in Niu (Fig 5.4.6 lag

F93-2A-11L3

3.4 I_,_rldin_ Load A torque the fuselage

Torque of 81,300

during

in-lbs

emergency

is imposed

on

landing.

buckling

load.

ultimate

margin

of safety

the members different

Tc=F_*d_+F_*dcwv F_=(IO0) (1.5)=15001bs Te=2700(21.5)÷1500(15.5) T¢=81,3OOinlbs

used to design

torque considerations,

thickness

and rivet spacing

withstand

the buckling

torsional

moment.

8,470 psi.

resulting

3.5 Fuselage

of safety

Bending

The weight

reacted

method,

through

psi.

of the engine

WeiEht

structure which

longerons for a

the engine

longerons

during an emergency must withstand

an 18g

rivet

on each rivet stresses

was 0.324

rivet

a 1/8" diameter spacing

stresses

of safety

on each rivet

was

and the ultimate margin

rivet

and assuming

of 1.0 inch, the maximum

on each

is 3.4 lbs.

The bearing

rivet is 281 lbs and

The ultimate

and bearing

is over 80.

Mount

The design mounting

is 510 lbs.

and the bearing

stress o12 each rivet

and ultimate

landing,

the

Skin

Using

margins

and assuming

was 0.263.

3.8 Engine impact

stress

429 lbs, respectively. of Engine

can be

can steer

of 1.5 inches,

and ultimate

Upper

bearing

at

at impact.

a 3/16" diameter

of safety

maximum each

of

Sldn

bearing

The bearing

is

Lon_erons Upon

fuse" that

direction

rivet spacing

maximum

3.7.2

of 1.56.

Considerations

a "mechanical

Using

of safety

applied a

The eleven

This technique

675 lbs and 966 Ibs, respectively,

a series of 15 inch C - section

of safety

Lower

margin

due to Engine

to the spars.

3.6 Bucldin_

stress

The

carry a stress of 15.6 ksi, thus allowing margin

is

The thickness

SpaeinE

a maximum

is 0.09.

stress on the fuselage

longerons

the critical

3.7. I

to

The stress due to shear flow

of the skin is 13,900

margin

bending

must be determined

loads caused by the

Using Niu's

for buckling

the skin

have an

to allow buckling

conditions.

in a desired

3.7 Rivet

of 2.14.

can be changed

loading

the engine

As with the engine

The C - sections

forward

Bolts of the engine

bolts to withstand

mount

requires

the shear load of an

a

F93-2A-IR3

l$g impact.

With

four 112" bolts the required shear

From

this

data

decomposed

load each must withstand is 10.8 ksi. A sheet

the

percentage

onto

the

of these

front

and rear

loads

were

spar.

thickness of 0.16 inches is necessary to avoid tear XA_

CM*C_ Cz.A

-0.07

(58.6) 1.4

=-2.93

x_-

CM*CCL=

-0.07

(58.6) 0.45

=-9.12

out around the bolts.

3.9

Nose

Gear

The

plate The

Landing

shear

and bolts

stress

was

supporting

Load

on

the

calculated

C - section

the

loads

through

the

the

plate

required

to avoid

nose

gear

using

four

is designed

fuselage.

The

mounting

Using

1/2"

bolts.

spar

carry

balance

with

of

pressure

was

to help

thickness

a tear out was

0.0951

inches.

3.10

Wing

The

position,

From

these

spar

carries

equilibrium

the the

on the

normal

force

of the

speed

while

wing

1093

the

rear

front

required

center

to

of

pounds.

shown

wing

the

that

the

loading

spar

front

at

carries

12%.

Reactions

fh'st task

undertaken

in the

design

Mps=

of

(Na *XA) - (Fas* (24.6)

=(9182.2.93) the

spar

at the

it was

total

about

rear

to be

calculations,

88%

analysis

force

calculated

maneuvering

Structure

3. I0. I Lift

a static

structure

was

determining

the

) =0

- (Fas* (24.6))

=0

FRs=IO931bs

loads

FFs=NA-FRs =9182-1093 induced

on the

wing

at the

maneuvering

and

dive

F_s= 8089

ibs

rear spar

bears

condition.

ra=CL*qa*S_ =1.4 (43) (152) =9150.41bs

D,= (C_o+kC_)*qA*Sw =(0.00785÷0.056"1.42 =7691bs

) (43)

NA--Vr'('9150.42 ÷7692 =91821bs

(152)

) At

dive

speed,

to the

rearward

at low

angles

the

front

spar

the

movement

of attack.

again

of the

Summing

yielded

the

a greater center

of pressure

moments

following:

load

about

due

F93-2A-

1R3

3.10.3

Mps= (_A*x_)- (Fro.(24.6)) =0 FRS=3404

Spar

Shear

Once

the

total

lift

the half-span load

distribution

condition

was

between

determined

the

spars

to be 63%

at the

for the

dive

diagrams

distribution

of the

were

and

for the

37%

rear.

Diagrams

sized

the

front

spar

condition

sized

the

rear spar.

was

the

determined

shear

versus

wing

and

for

moment

station.

front

So, the maneuvering

condition

wing,

plotted

WING spar

Moment

ibs

Fes=N A -F_ Fes=577 81bs

The

and

SH_A_

SHEA_q

DIAGRAM

ILO)

and the dive 18

Sl

3.10.2

Total

Wing

Lift

Distribution

,IQ

20

The

external

lift distribution

on the wing 4!

2|

structure

was

trapezoidal

lift

the

two

gave

lift

distribution

shown

determined

using

distribution.

a close

Taking

the

approximation

of the

m figure

an elliptical

wing.

This

average

to the

II|

14|

I_S

II0

Ill 152

WIN(]STATIONIIIIl

Figure

of

actual

lift distribution

I|

I!

and

2

WING MOMENT is

12

II

1.

I

IIOgEN1

[II_,L

IS1

DIAGRAM

([llllllll)

\ \

i

\

4

,....

?

20

40

.TJ nU.

I|

I|

1O0

I21

141

110

Ill 152

WIN6 STATIONItN) J

Figure

3

• .'_._-_.

% % I

_4

I N

-" ---

Figure

l _

I m

VmiNma_I._ Ibmmbwmm ^'mmqp t,,_ Dimm'bxtlou

1

I so

I ml

I lm

I ill

I iI

I m

Using

the

percentages

shear

and

moment

were

produced.

calculated

diagrams

m section

for the

respective

3. I0.1,

spars

F93-2A-1R3

FRONT SHEAP IDQ

SPAR

SHEAR

maneuver,

DIAGRAM

(t95}

x, \

and flapped

aerodynamic

change

the following

relation,

conditions.

The

in torque was calculated

using

BO

& T=-q*C.*SLoc*c_oc AT=(119.5) (-0.07) (3.07)

4_

A T =--74.73

--...,

(2.91)

inlbs

"%

?!

20

i|

40

II

tl!

IZ0

140

1|0

It!

52

over the various

WlllG STATIOg (111)

Figure

4

produced

flight condition.

The boom torque,

by the tail force, was determined

as

follows, FRONT NONENI

SPAR lINt

B$I

MOMENT

DIAGRAM

[lillSJllt)

nz _w

f4 .4 210--95_ ffi6 0 .51b/ f C

\ \ \ From

Fig A5-FAR

23 Appendix

A

w=381b/fC 2 Lr=w*Sa.zffi (38) (55.3) |1

1[I

II

tO

100

120

140

}SD

IIlO

152

Lr=21OOlbs

W iilG STATION (IWI

Mz=Lz*d_= (2100) (165) M_=346500inlbs

Figure

5

From these

diagrams,

The landing the attachment

are determined s/zed

loads are also present

at the boom

lugs were to be,

based on the Joads at the wing-fuselage Fzfp*Wac*3g=(O.8) (2091) (3) Ff=5Olelbs TG=Ff*dv+Wac* (3g) *d h TG=(5018) (38)+(2091) (3) (21.5) TG=325600ibs

interface. 3.10.4

Win2

Torsion

The wing torsion span of the wing landing

loads.

aerodynamic

through

accumulates

over the

aerodynamic

loads and

The torque produced loads were calculated

by for dive,

These torques produce

are plotted

the torque

versus wing station to

curve shown below.

and

F93-2A-1R3

REAR

SHEAR

SPAR

DIAGRAM

SH EAI_ [L85)

\ \ \ From the torque curve, the rib and stringer spacings OI 0

i M

i|

4e

|n WIHG

Figure

I1|

IZl

STATION

141

I|1

110

determined

characteristics

of the skin panels.

3.11 Environmental 3.11.1

SPAR MOMENT

Temperature

allow

\

occupant

conditions.

'\

"x. _ e ZO

i _0

|a

l|

1OO

WING

Figure

ventilation

has been incorporated in high temperature heating

provide

warm air to the cabin

weather

conditions.

140

Ill

joints

rio IS2

are designed

for thermal

All structural members

3.11.2

taking

into account

expansion

. ,

.:,.::,_:::

:::

range(-40degree

Atmospheric

:: :: : :: ::::-m

requirements pressurized States.

F to +122degree

F)

Pressure defines

the

for supplemental civil aircraft

oxygen

registered

The cabin of the aircraft

in non-

in the United shall be equipped

_ 41

61

81

101

121

141

1G1

181

w_asmmon) Figure

tolerances

:::]o100

:iii:iiiiii:iii: iii:iiii= • tBL 21

and

for the given thermal

FAR Part 91.211(a) :

in cold

of the aircraft.

Wing Torsion

_:

for operation

will

IINI

7

:::::::

system

.-......,...

II0

5TATIOH

comfort

A supplemental

operating

-10!

to

II|o,l|llsl

,,xN

01 |

Considerations

DIAGRAM Cabin

IIX.tISI

WOW_NI

using buckling

_!2

[INI

6

REAR

were

with supplemental

oxygen

for necessary

crew

8 members 9

and passengers

for the given conditions,

F93-2A-1R3 up to

16,000

3.11.3

Sand

feet, defined by the FAR's.

humidity

and Dust

windshield

All external

surfaces

will be coated

chip resistant

aircraft paint to withstand

sand damage

specified

will be outfitted

in the SOW.

with plastic

moving

parts to reduce friction

particle

matter.

All windows

rubber weatherstrip

to keep out excessive

Air filters

are to be installed

ducts

to prevent

engine

system 3.11.4

damage

3.11.5

to water

from 2024 aircraft under certain

sand and

washers

between

prevent

icing

control

3.2.7

the conditions cabin

metal

have

inches

of wet snow.

Filters

are

accumulation.

Drainage

in the air ducts

at locations

from 2024

skin panels aluminum

temperatures

in the SOW.

and pressures,

3.11.9

At

Wind

between

to withstand I0

by the weight

of 20

are manufactured

to resist

and interface

corrosion materials

moving

corrosion

door hinge

also. parts

as well.

and Gust

The fuselage

excess

loads m

that will resist corrosion

will help to prevent

and resist corrosion

specified

washers

supporting

to support

aircraft aluminum

have been chosen

are manufactured

and interaal

skin panels

from saltand fog. Rivets

Plastic

or plastic

Atmosphere

All external

and icing.

Teflon

hinges.

encountered

3. I 1.8 Salt/Fog

ducts to

and subsequent

parts will help

been designed

and/or sealant

to prevent

with seals

Snow

of those

accumulation

hinge

of the door

excess

damage.

inmasion

of the latch mechanism.

Humidity All external

water

and doors will be outfitted

water

are to be made

snsceptable

to prevent

All skin panels

in the engine air intake

excessive

condensation.

freezing

in the intake

defrost

into the aircraft to

with

and environmental

and subsequent

to be installed

holes

designed

members

intrusion

A heated

All doors are to be equipped

Rain

with weatherstrippmg

prevent

windshield

on the front

Ie..._e

by sand and dust.

All windows

water

3.11.6

caused by

will be sealed

and side windows.

eliminate

hinges

to condense

has been integrated

between

damage

dust.

system

dust and

Door

washers

with

is expected

structure

gust loadings

has been designed

as def'med by FAR Part

4--"

F93-2A-1R3

23.

All tie-down

wings 4.

will be attached

to the f= M_y= I

and tail structure.

Structural

4.1

fittings

Floor

(727200) (8. O) 198.22 f=29.3ksi

Substantiation

Structure

Bending

Under aerodynamic

Fry = 47 PSi

0vlS)ty

Ftu = 62 Psi

(MS)m=

= 0.60

Due to Occupants loads,

0.41

the occupants 4.3 Skin Bucklin E Due to Landing

exert a bending

moment

on the floor members

The torque needs

to be counteracted.

during 7075-T6 Values

the landing

= 2.9032

(82800) (I.5) 2.9032 f=42.8ksi

by the torque

Fty = 76 ksi

(MS)ty

= 0.78

Ftu = 83 PSi

(MS)tu

= 0.29

The torque

This is a worst with the entire weight bending

moment

4.4 Skin Bueldin_ The torque

longerons model

and underfloor

yields

Due to Nose fear

a I total=

(see dwg F93-2A-102-7

Load

the fuselage

ease of the aircraft landing on the nose wheel.

is transferred

through

members. 198.22 ; Parts

by the under

(see dwg F93-2A-102-7;

Section

produced

is 81,300

inlbs through

skin.

that is produced the

engine

The calculation

f=_=_=136p.s.i.

in _ and y = 8.0 in

Fcrle=222p. S. i.

1,2,3). (IvlS)bu = 0.09

is the same as

changes,

re 12000 q=_= (2) (1773) q=3. 401b/inch

fuselage

by the engine

ring frame, and into

This

11

in.

Due to EnLwine Torque

that in 4.3 with the following

The simplified

of 0.040

= 0.09

is reacted through Bendin2

box formed

an area of 120 in2 and skin thickness (MS)bu

Fuselage

stated in FAR 23 is

in 4. A-A).

4.2

criteria

for the beams floor carry through

in 4 (4 beams)

f_My_ I

produce

channels absorbed

are Ixx = 0.7258

that the occupants

This is accomplished

with the use of 4 NAS346-45 (see dwg. F93-2A-102-7).

Load

that

e.-

F93-2A-1R3

4.5

Fuselage

Bendin_

The moment

Due

to Ent./ne

produced

Load

V_

(5961)

fs-A

by an acceleration

(7)

(1.5) (1,2) (_)

(4)

fs=13.7ksi on the engine is reacted through

11 longerons

fs _ V_

(5961) (1.5) (1.2) 1 (_) (0.1) (4)

DC spaced

around

the periphery

dwg F93-2A-102-7;

of the fuselage

(see

fs=53.6ksi

The stress in each

Part 15).

Fsu = 75 ksi member

(lVIS)su = 2.64

is, Fbu = 118 ksi

f=Mmr..Y=

(83072) (1.5) =15.5ksi 7.99

4.8 Engine

(MS)bu

Mount

Bolt Sizin_

The engine Fty = 47 ksi

(IVlS)ty = 2.0

Ftu = 62 ksi

(MS)tu

with

Seat Track

is connected

four gusset plates.

These

to the ring plates

frame

allow

the use

= 1.6 of ANgC-7

4.6

= 0.46

bolts,

one per plate,

to transfer

the

Fasteners engine loads

The seat tracks

are fixed

ttzrough

into the ring frame

and longerons

(see

the floor dwg F93-2A-102-7).

panel

into the NAS346-45

channel

with four 5/16 V_

in screws

(8496) (1.5) (i.2)

per track. fs-A

(_)

(

)2(4)

fs=19.5ksi fsfV__

(6210) (1.5) (1.2)

f_

v_ DC

(_) (0;I) (4) fs=47.8ksi

fs=36.4ksi f_

V_ DC

(8496) (1.s) (1.2) 1

(6210) (1.5) (1.2) (_6)

(.188) (4)

Fsu = 75 ksi

(MS)su = 1.56

Fbu = 118 ksi

(MS)bu

fs=47.6ksi

Fsu = 75 ksi

(MS)su

= 0.05

Fbu = 251 ksi (for steel)

(MS)bu

= 5.27

4.9

Spar

Cap Sizing The principle

4.7

Nose Gear

of the section

modulus

was

Bolt Sizing used to generate

The nose gear is held

the spar cap sizing.

with 4 ASN$C-11

The caps start

onto the forward off as an NAS344-69

bulkhead

= 0.64

for the front

spar and -33 for

bolts (see dwg F93-2Athe rear spar (see dwg F93-2A-104-7).

The

102-7). justification 12

for the sizes is detailed

below.

a-

F93-2A-1R3

4.10

Spar

Ca.

Fati_e

s= I---z=z" y °i (_-_=) Fatigue SR_=I.

01

( 88282 78000

) =5.03inches "

s,_ZL = _11.55

=

modulus

wing for the both spars yields

over the spars as designated

the following

Pfs=920

lbs and Prs=3241bs

inches.

Using

ZI

12

12

WING

the S-N curve

Figure

from Fig 3.7.4.1.8(g)

and the following,

(940) (44) S (5.63) free.a=7190p, s. i. fmx=2.2 fme_=2.2 (7190 ) fmax=1518p. S. i. N=2 (10 s) cycles

I1|

Ill

14!

Ill

III

the rear spar yields

a count of 5(10 _)

cycles.

l!2

5TATIOW [IN]

r1211111

--S

_S

4.11 Shear

111i11211

Web

Sizin2

9 Shear

flow is determined

diagram and the web thicknesses

SECTION

MODULUS

REAR SECTION

IIQOVLU$

meeting

the budding

criteria.

from the torque are justified

A sample

by

calculation

SPAR is shown below.

I_N )

2.5

2

of 44

f_Pl_

SPAR

[IN :=)

I0

at a distance

curves.

Similarly,

2

in 3.10.1.

MODULUS

FRONT IiODIIHI$

of the aircraft

3

from MIL-HDBK-SE

5FCIIO)I

on the

over the span of the

plotted

SECTION

were performed

spar caps using half of the weight

5 .63 inches

distributed The section

estimations

3

.',,, q=7 09 ib/ in ch

1,_ h

t

fc._.=5672p. s. i. f_c=fc.c.(i. 5) =5672 (I. 5) f_zf8508p.s, i.

2,5 _'1 _

-------

..-...

,.__....i,__.

'

0 2

_2

42

II

li

WIWG _5

IIllllll

101

I:'J

120

122

_111 IS2

STATIOW {IWI ",-,I--

F=_t=_

_ Illnl1211

" _,2

0.125)2 J =5 (107)(-_T_-

F=zlt=10,088p.

Figure

S. i.

10 (MS)b = 0.19 (see dwg F93-2A-104-7

13

for remaining

thicknesses).

F93-2A-1R3

4.12

Rib

Sizin2 Kz_ From

the

torque

diagram

the

value

fs torque

then

at WS172

sized

can

using

the

be

determined.

buckling

criteria

The

s-d= s

for

rib

5.5-. 375 5.5

=__g_ _ 709 t,,_= Kr _ (0.932) (0.125)

See

drawing

used

F93-2A-104-7

=--q (___hD )

for a 709 fs_,o- 0 . 125

complete

list

of thicknesses.

q= 1.76 f

= q= z,..r,. _

fs_

4.13

ib/inch

Lug

(

8.8 8 .8-5 .25

=14.

lksi

Justification

1.76 .--_=55.1p.s.i. The

f_=fL.L.

=6. Iksi

is

f previously.

=0.932

lug's

purpose

is to transfer

shear

and

(1.5)=55.1(1.5)=82.7p.s.i.

f=_ t=KE (_ )2= 7

(10 T) ( "032)2=179p.

S._

--fff-

bending

They

moments

are

to the

subject

fuselage

to the

highest

carry-through.

shear

and

bonding

(MS)b = 1.2 moment 4.12

LiEhtenin_

Hole

4.13. Lightening

since

they

holes

were

considered

inboard

of the

I Forward

boom

due

to there

thickness

Using

the

calculations

in Niu

and

the

following

dimensions

were

inches,

5=7.7

Lug

inches.

Values

q=709

lb/in,

rivet

to the

for the front

88282

shear

and moment

spar

the

values

are

919.6

Ibs

inlbs

at the

root,

The

lugs

respectively.

5.8

inches

apart which

yields

a loading,

P, of

previously 15,200

determined;

wing.

obtained; are

D=-5.25

of the

figure and

6.2.3,

root

and diagram

weight.

at the

for the According

ribs

are

Justification

diameter=.375

lbs.

The

bolt

is in quad

shear

so the

stress

in., is,

and rivet

spacing=5.5

in.

fa-

P (4 (A)

fs- 15,200(1.5) 709 fs_,_- 0. 125 fs

(

7.7 ) 7 .7 -5 •25

(1.2)

(_) (.75

(4.4)

(4)

fs=68ksi

=17.Sksi

(Fsu)bolt

The

steel

14

0vfS)su

= 75 ksi

material

because

for the lug

was

chosen

of its high

strength

= 0.1

to be 4340

properties.

After

at-

F93-2A-1R3

calculating

flange

for bearing, thickest flange

thicknesses

tensile,

the tensile

came from tensile

stress requirement.

Rear

stress.

longeronsof the Quest PFT are manufactured from The

7075-T6 Aluminum

to satisfy assessments

Aluminum

life of 107 cycles.

the bolt diameter

The entirefuselageskinisrivetedtogetherusing

as the front lug

was found to be 7/16 inches

the flange

thickness

As before

the lug thickness

and

in quad shear was .114 inches.

of stress and tensile

was checked

for each

stress was the sizing

Aircraft

use AN430DD-8

rivets.

designation

bolts are used to mount

type

gear.

changed

under

the belly

account

for the torsional

quality

AN the engine

The skin thickness

was

of the aircraft in order loading

produced

to

by the

occupants. Carry

Through

The carry designed

through

are taken

structural members. from the moment moment

of inertia

5.2 Enmine

Structure structure

to carry a pure bending

shear reactions

for the wing

moment.

is

The

attached brake

up by the fuselage

The pure moment, diagram,

determined

Irear_10.8

reinforcing

in4) for the

and Maintenance

Engine

Provisions

brackets

maintenance

of four main

5.3 Control structure

assembly

components:

aluminum

to a ring frame

bulkhead meet

by the use of

e channels.

The

to the ring frame

with

attached

to the longerons.

is accomplished

cowlings

located

through

is

System

and Ventilation

The control system

longerons,

requires 15

allowing

space

the

on the left and right

side of the engine.

Assembly

The fuselage comprised

formed

use of engine

Manufaeturin_

5.1 Fuselage

to the rearward

engine truss is mounted

is taken up in the

(Ifronr=31

Mounts

The engine mounts

carry through. 5.0

rivets, while the internalcomponents

and nose landing

factor. 4.14

sheetbrake formed and fittedto shape.

AN456DD-4

the same procedure

extrusions, while the remaining

components aremanufactured from 2024-T3

Lu_

Using

type

Fatigue

ring frames,bulkheads,and skin. The floor

the

needs to be .125 inches

were made at an infinite 4.13.2

over stresses

tear out, and fatigue,

requirement thickness

required

installation

Routin_ simply

for the routing of this

F93-2A- 1R3

system.

This is provided

via a channel

between

of the four floor

longerons.

will incorporate

push-pull

to the control

systems

plates

under

located

system

rods for control.

are provided

the two

located

under the engine

via the air cooling is easily

of the engine 5.4 Front

vents

Access

is

section

properties.

from 2024-T3

on the engine.

using two 2024

This

from the same location

shape with varying shear quality

flow.

These

NAS

5.5 Front

T-sections

The shear

Aluminum

1304-4P

and Rear

shaped

composed

to the designed

of a high

strength

to the moment

the wing.

The rear spar was modified

for the

a half-hoop

aluminum

spacing

c-section

Wing

location,

support

while

on

Structure

and

Skin

of the ribs, along with the from the torsion

induced

ribs are manufactured

from 2024

and fitted

to shape.

were also placed buckling

at the

for the engine.

loads of the aircraft.

the required 16

c-section

This was done in order to

were designed

stiffeners

steel),

sheet

slightly,

to a straight

for the engine

formed

sized

at the root of

from the aerodynamic

brake

It is

formed

by 24 bolts

produced

was attached

The sizing

of one

configuration. steel (4340

according

5.7 Internal

Lugs

brake

to the firewall

same time providing

using high

sheet

of the

was assembled

attached

compensate

steel bolts and rivets.

is comprised

aluminum

was one

in the design

spar carrythrough

top of the hoop.

formed and fitted to

structure

c-sections

brake formed

with varying

to account

diameter

Structure

encountered

The front

wherein

from 7075-

are connected

The lug assembly piece

of

web is manufactured

sheet

thickness parts

Carr_tthrou_h

of the difficulties

the spar caps and shear

extrusions

spars with the use of 0.25"

wing.

The spar caps are manufactured

to the spar cap on

The wing carrythrough

Spar

two main components:

T6 Aluminum

5.6 Winz

in

The blower

to resist

bIAS bolts.

The venting

The front and rear spar is comprised

web.

the respective

doors.

and Rear

plated

The lug is connected

and fresh air is provided

maintained

cowling

may have to be cadmium

corrosion.

by inspection

the fuselage.

side longerons.

which

This system

runs along the side of the aircraft

between

system

running

aircraft

aluminum

Various

in the wing

criteria.

The

to help meet

The entire

wing skin

F93-2A-1R3 is rivetedtogether usingqualityAN is manufactured

from 2024

0.032" thickness

outboard

0.05 _ thickness was designed

inboard

Finally

were

at strategic

placed

provide

aluminum

various

torsion

access

and

produced

Weight

by

Thereby,

and control

passenger

design

of the aircraft is an essential

aspect that must be addressed

maximum loading providing savings

stiffness conditions, a light,

while

Weight

by using aluminum

cost. Additional

punching

lightening

weight

holes in the extruded

Extruded

than brake formed

savings

parts are usually

members

Rivet weight them

weight.

The total weight

ratio include aluminum thicker

so they tend to be

is compensated

allowing

sheet

since it

high stiffness-to-weight

at a low

heavier.

at the same time

of the structural members

a relatively

floor longerons.

for given applied

yet durable structure.

are accomplished

for the majority provides

of members

to allow for

to add 1% to overall

for by structural

for the aircraft structure is 17

weight

an increase seems

seater

in 13 lbs

the loadings

seems

of the aircraft wing

for the two

in the two

An increase

to support

by a third passenger

The weight

to

for the structure

was 131 lbs.

The total weight

Summary

The weight

version

imposed

systems. 6.0

The weight

for the structural

or holes

in order

of the wing

seater

This

panels

locations

144 lbs.

sheet of

of the tail booms.

for maintenance

the skin

of the tail booms,

due to the large

the booms.

rivets,

version

to be valid. is 269 Ibs. was 253 Ibs.

of 13 lbs, also, for the third

to be valid.

z-

F93-2A-1R3

wm 8 Weight

Estimation

PART

QTY

GAGE

WEIGHT

(IN)

O_B)

RIB WS 192

2

0.016

0.701344

RiB WS 172

2

0.016

0.788608

RIB WS 148

2

0.032

1.881024

RIB WS 99

2

0.032

2.495104

RIB WS 87

2

0.032

2.676096

RIB WS 75

2

0.032

2.805376

RIB WS 49

2

0.05

4.9288

RIB WS 42

2

0.125

10.12525

RIB WS 35

2

0.125

10.12525

RIB WS 28

2

0.125

10.12525

RIB WS 21

2

0.125

10.12525

F-SPAR

CAP 0-80

2

12.84256

F-SPAR

CAP 80-120

2

11635696

F-SPAR

CAP

120-192

2

1.6704

R-SPAR

CAP

0-20

2

1.102868

R-SPAR

CAP 20-100

2

3.464398

R-SPAR

CAP

2

0.531468

100-140

BUCKLE

STIFFENER

(#9)

44

0.02

3.26634

BUCKLE

STIFFENER

(#10)

28

0.02

4.010104

BUCKLE

STIFFENER

(#11)

12

0.02

2.63004

BUCKLE

STIFFENER

(# 12)

4

0.02

1.25846

SHR WEB

(20-49)

4

0.125

12.4432

SPAR SHR WEB

(49-76)

4

0.05

4.52682

SPAR SHR WEB

(76-172)

4

0.032

7.632691

SPAR SHR WEB

(172-192)

4

0.016

0.58176

SPAR

18

F93-2A1R3

FRONTLUG

4

1

12.9614

REARLUG

4

I

6.9618

FRONTCARRYTHROUGH

I

0.09

11.69883

REAR

1

0.05

4.70862

CARRY

THROUGH

SKINS:

0

WING

TIP

OUTBOARD MID

LEADING

EDGE

MID

2

0.032

1.105344

2

0.032

16.65126

2

0.032

18.68096

2

0.032

30.75571

INBOARD

LEADING

EDGE

2

0.05

31.5524

INBOARD

TRAILING

EDGE

2

0.05

12.221

INBOARD

BOTTOM

2

0.05

7.373

WEIGHT=

269.0445

LBS

7.0 Conclusion

aerodynamic

The goal provide along

a cabin fuselage with an adequate

the loading

criterion

Control systems designed Volumetric assured

wing

defined

detail

is to

components

design

landing

decomposed

by FAR Part 23.

structure.

routing

cabin requirements

were also met that

spacing

Calculations

induced

of the torsion,

floor design these

required

in which loads.

a different

produced

a change

Calculations

The location

carrythrough

or shear

stiffeners

into the sizing This included

required

and safety of the

by the occupants

to counteract

parts.

by various

inboard

sizing the front

approach

and rear

Also the large

of the tail booms

were

torque.

The cabin was expanded

of the

three-seat

the wing

torsion

required

to counteract

configuration.

other

on the mid-fuselage

to designing

to the skin thickness

passenger

and

skin, ribs, along with

of the wing

structure.

performed

of the total wing

change

19

aircraft

such as the twin tail booms, the

spar and lug, the wing

have been

and removal.

induced

gear, and the wing were

structure that meets

for ease of maintenance

to the original added

project

structure

and ventilation

adequate

occupants. flow,

of this design

forces

a this

to accommodate The new engine

a

F93-2A-1P.3 location

required

providing

account

for a longer

and sizes provided

moment

a ring-frame arm. Weight

by the preliminary

structure to estimates

design

report

were for a two-seat

configuration

assumptions

to be made about the weight

this larger Finally,

needed

of

aircraft that will need to be validated.

environmental

for all relevant requirements, performed

aircraft. Certain

parts.

conditions This design

were addressed meets all of the

but further optimization

through

may be

later iterations.

2O

Appendices

.t---

lOO0

--

-

loo

|

!

I I _ i II

1

_._

_

_

,

-

_E

I

|

| d _ |_

_ Y

Theoretical

I

- .

I' h I _111

A

"t '

10

"" 7

f

t

t

ltlfl

t

I

Irlfl

70

I 100

I IIIIII 1000

I !

Fig. 5.4.5

Compression buckling coefficientsKc(circular cylinders).

12

10

gl

8

3

4

5

6

7

_(a - tonsside)

Fig.5.4.6 Shear buckling coefficientsKs (circular cylinders). Airframe P_

PAGE

Structural

Design

139

I_'LA.!4"K NO]" FP.J,,'.,_::D

1 6.2

Lightly

Loaded

Beams

D _< -- < 0.75; h

0.25

The ideal construction for most shear-carrying beams is a tension field (or diagonal tension beam per Ref. 6.8). However, in some cases it is advantageous, and in other cases necessary, to incorporate circular, flanged holes in the beam webs. These cases come under two main categories: • Lightly loaded or very shallow beams. In such eases it may not be practical to construct an efficiently designed tension field beam because of minimum gage considerations and other restrictions due to the small size of the parts involved. It may then be advantageous from a weight standpoint to omit web stiffeners and, instead, introduce a series of standard flanged lightening holes, as shown in Fig. 6.2.1. • Moderately loaded beams with access holes. Where it is necessary to introduce access holes into the web of a shear-carrying beam, a light, low cost construction is obtained by using a flanged hole with web stiffeners between the holes.

web thickness

0.016

_< t _< 0.125

0.3 _<--D_< 0.7: b

40 _<-h _< 250 t

H

_

Cap

centroid

. /

.

-

-Z--+_ o

"[

Type I d -_--

rivet

rivet

diameter

centerline

Lightly Loaded or Very Shallow Beams The following two types of beam construction are considered. The standard flanged lightening holes as shown in fig. 6.2.2 are centered and equally spaced. • The limiting conditions for the design curves is given in Fig. 6.2.3.

i D

H

i

2.0 2.5 3.0 3.514.0

4.5 5.0!6.0

{.Vote:_

[

is

Fig. 6.2.2 I (inch) ,.25 .t .3 , .4 !i.45 J .5 i .5 i'.55 f (inch) '.5 i (a) Lightening holes of typical

the

assumed

effective

depth

of beam

cap)

Lightly loaded or very shallow beams.

_j_: I

I, _.L__.I_ ;t

flanged)

r.

II

I

2Si

,.-K.,. 0._

[

_.

I

O.-- Outside.diameter - o.lss i.o.I

22 _-

Do (Inch) 1.7 1.95 2.65 3.0 3.65 3.9 4.95 5.95 6.95 7.44 7.95 8.95 9.45 Fig. 6.2.1

D-

Inside

D (Inch) 0.8 1.05 1.7 2.05 2.7 2.95 3.8 4.8 5.8 6.3 6.8 7.8 8.3

diameter

a (Inch) 0.2 0.2 0.25 0.25 0.25 0.25 0.4 0.4 0.4 0.4 0.4 0.4 0.4

Common flanged lightening (t - 0.032 in -- 0.125 in)

,,

-\

o.

18

\,

-,

_

I

0.: o.

t I

I I

" 10

holes.

,q

2o

14

(_ L_enmg holes with bcadc.d

.

\1 \1 ,,N

"l£-4. ',,,, \

_ o.-_

/

0.5

\\Na\\

l I

I\"( 0:s5 t \07

'

0.2 o.-_ o., o.s o.6 0.7 i 7075-T6

,=o 6o _o _oo _

_,=o_

f

fig. 6.2.3

;

"

_so zoo 220 _o

t

Ultimate allowable,gross shear stress for aluminum alloy webs with flanged holes as shown in Fig. 6.2.1(a).

J ....

1 !

Airframe

Structural

:r'L:_r_'.ll3_r'_ PA_,E BLANK

Design

NOT

165

FM.MED

1 November

1990

_=_.|

70, ...... : : : :

: : : :

: : : :

: : : :

......

i

": _ .'-"-_

i _

i i

i iiii i i_i

.......

: : : : :

:

i i i

:

:- _ _ ..-..-_:

i i ':

i i iii_ _ !i'gi .: " : :'i

:

_ -""-: "

: : : : :

:: :: :: ::

:

: .

i

: : : : :

: : : : :

:

:

: : :

i i i

i i iiig _ iiVi i i!iii

:

:

:

:: .-: :: :: : :

: :1.

i ..

i ii!_i :. _ :- . :..:

:

:

i

i i iiii

:

_ _ i li ii

:

:

!

_x .:

i i iliii

:

:

:

:

:

:.

-

:..._.

i

!

i i i i!!

FIGURE 3.7.4. 1.8(g).

Properties:

.TYS, ksi

Information

RT

Details:

.

-: _ - -: :

: :

:

:

: :_"

:

:

:

:

:

: :::

: :

:

:: ::

;

:

:

: :::

i

:

- . ...:

_

i

i+iiii

: :

: :

: :

: :

: :

: :

ii

: ::: : :::

:: : :: : ::

Notch Type

Gross Width

Net Width

Notch Radius

Edge Edge Fillet

2.25 4.10 2.25

1.500 1.500 1.500

0.057 0.070 0.0195

ill!i!

C¥C_.S

K, = 4.0, 7075-T6

aluminum

alloy sheet,

for Figure 3.7.4.1.8(g)

No. of Heats/Lots:

Notched

Equivalent

Not specified

Stress Equation:

Log Nt - 10.2-4.63 log (S_q - 5.3) Seq = Sma_(l-R) °'_ Standard Error of Estimate = 0.51 Standard Deviation in Life = 1.08 R 2 = 78% Sample

Etectropolished

3.2.3.1.8(b),

_ i i_.:.

:

Loading - Axial Frequency - 1100 to 1800 cpm Temperature - RT Environment - Air

(unnotehed) RT (notched)

Specimen

Reference:

!

Test Parameters:

Temp., F

76

82

Condition:

i i!ii

::

Best-fit S/ N curves for notched, longitudinal direction.

82

Surface

LII_,

Bare sheet, 0.090 inch

TUS, ksi

i

.....

Correlative Form:

:::

:::

FFITIGUE

Product

:

.......

: :

:

:

.............

i

Size = 126

[Caution: The equivalent stress model may provideunrealistic life predictions for stress ratios beyond those represented above]

(f), (g), and (h)

3-366 . .,(,,_¢_',o;,_ v,._';_

I_ #rcK

NOT

FtL _,!"___

MIL-HDBK-SE 1 June 1987

°41 1iliiiilt iiiiii!i .

++ .........................................

•

_ i _i_ .+

)

i i l)i))

.....iiiii

:

:

:

:

: ........

:

:

: ::::

i

i

::::

.....

i i ili)_

:

:

:

:

: :

:

:

:

: :

: : ::

: : : :

: : : :

i: i: iii : ::

i:

_: i: ili : ::

: :

: :

: :

: : : :

: : : :

: : : :

: : : :

:: :: :: ::

: :

: :

: :

:: :::

!:

i ii iiii

: :

......

.........

i

: :: : ::::

: :

: : :: ::::

...... _

-

; _-_

:

:

:

: ::

:.,: : " _I-_

:

:

:

-

:

:

.... : :

: :

: :

:

:

:

, .

-

"

: ! "..

_

_

"_ ! _

:

:

:

:

:

:

: ::

:

:

:

: :

:

i

" _ !::

_

"

-

_ -i

i

) )i)ii

°" i

:

:

: :::: : ::

......!i i :::::_

io+"

) ++)i+)+"+ ))+)+

! i i!+ii

O,

"

:

+

:

:

:

::*.

:

.

:

: ::

_o' FIGURE

3.2.3.1.8(ta).

Best-fit S/ N curves for longitudinal direction. Correlative

Product

Form:

Bare

sheet,

0.090

notched,

Information

TUS,

ksi

TYS,

73

Temp.,

54

67

--

RT

Notch

Gross

Net

Notch

Type

Width

Width

Radius

Edge Edge Fillet

2.25 4.I0 2.25

1.50 1.50 1.50

0.057 0.070 0.0195

Surface

Reference:

Condition:

3.2.3.1.8(b),

sheet,

RT -_ Air

No. of Heats/Lots:

Equivalent Notched,

alloy

Axial - l I00 to IgO0 cpm

TemperatureEnvironment

(notched, K,: 4.0) Details:

aluminum

3.2.3.1.8(h)

LoadingFrequency

F

(unnotched) RT

Specimen

io'

Test Parameters:

inch

ksi

K, = 4.0 of 2024-23

for Figure

1

Properties:

: ::: :::: :::: :::.

.+ . _.:.:-

))

......

: :

: : : :

i i i iilii

............

,.+ ..- ..:+.

: : : :

Stress

Not

specified

Equation:

Kt = 2.0 Log Nr = 8.3-3.30 log (S_q - 8.5) Seq : S_ (I-R)TM Standard Error of Estimate = 0.39 Standard Deviation in Life = 1.24 R-':

90%

Sample

Electropolished, machined, and burrs removed with fine crocus cloth

(e_ (f), (g), and (h}

Size : 126

[Caution: The equivalent stress model may pro_,'ide unrealistic life predictions for stress ratios beyond those represented above]

3-110

P_'-"EGIDIN_

PAGE

BLANK

NOT

FILMEIP

MIL-HDBK-5F 1 November 1990

220 STRESS

280

o :

180

..... .

168

:

÷

:

_

÷

: : ::::

" " • ":" "'.' '7

......................

:

! ":':'H

0.54

_IS!

:::::

_qTSO

-4.00

4342

R_N-C!JT

.... i i _! _!!! ....... i i i _ i!_-i_,,: ....! i _/!__!i...... ! _ i _: i _ "..... _i

v

.

K_-2.O

i '._":

_t4e m w n."

!

_- 120 D I2 100 I:

i i i!ii

i

i

: :

:::

: ",4..,: : :1::

60

..... i...i-- i-i- .i-i-ii._..... i... !._'i\

48

..... ii..... iii..... ,li........ .....................

i/

._ _!.:.i i

I0 s

2.3.1.3.8(n). Best-fit S/ N curves for longitudinal direction. Correlative Form:

LIFE,

notched,

Information

ksi

TYS,

266

ksi

232

390

--

Surface

Reference:

Details:

No.

V-Groove, K t = 2.0 gross diameter net diameter root radius, r

60 ° flank

angle,

Lathe

turned

4340

alloy

steel

bar, 1:,, = 260 ksL

2.3.1.3.8(n) Parameters:

cpm

1

Stress

Equation:

Log Nf = 9.46-2.65 log (St_- 50.0) Seq = Sraax (1-R)0.64 Standard Error of Estimate = 0.22 Standard Deviation in Life = 0.34 R 2 = 58%

to to RMS

of Heats/Lots:

Equivalent

Notched, 0.300-inch 0.220-inch 0.030-inch

Condition:

!_ _

RT (unnotched) RT (notched)

Specimen

_" ;-

LoadingAxial Frequency - 2000 to 2500 TemperatureRT Atmosphere - Air

F

:.-:

!0'

K, = 2.0, AISI

Test

Temp.,

..:.....

CYCLES

for Figure

Rolled bar, I-1/8 inches diameter, air melted

TUS,

ii

10 _

4

Properties:

..... i- -:'i"_'_-!

ii il F°_-__!

IO"

"

'.....:_

_',i',',ii',', 1i ii!iiiil: :,:...... i ..... :

FATIGUE

Product

STRESSES ARE BASED

80

28 103

FIGURE

NOTE:

.... __i_!:?_[.....?"i_::?il ....1iii"_4 .....i:ii!:i

m

x n-

!

Sample

10

Size = 30

[Caution: The equivalent stress model may provide unrealistic life predictions for stress ratios beyond those represented above]

2.3.1.3.8(a)

2-50

p_

P_GE

_t _.f_. NOT FSL_TC

,E,w+ :, _,'_

_.,,._ c

C-

_{ -ooe_

- o, tg

( ,:bo I.v/,_.-. 9go I_,I

,

...

IS"_O

=

3-<'0

_ ,SO

÷

_j_j

v,,'_.I =. gO

I_,_]

)_co = g_'$" lloo

"0,11_

=

_3g"

I0=o=

9_8

J =/0o

I= 2./_9

.2../o0

=.Z08'7'

..ze'oO = .Z_'Z. - 0,1',_

o.6_g2 -,2,_l# Iv,,,

/s_

!

I

I

!

'

.,-_£_.

.

for

n,_

"

z_

_=

'

534(n=_)

for

50

"

Q-

(2)

_= 4.8

-4-,534

< 47and

'

"

nl_"

>

. W.

"

I i

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I

I

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surface

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loadtn_

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MANEUVERING Fzou-_

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$0 WING Umlt

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60

(n,_)

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f%

(i)

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[

PART

I10

LOADING control

suttee

nl_ loading.

I00 POUNDS/SQ.

FT.

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FRAME

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20

z_O

60

80

0

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Contro

Ul

#OLDOUT

:ation 100

A

(in.) 120

I,

FPJ_,_

140

160

180

200

IIl,l,i,l,l,lll Leading Fwd of

edge front

Access spar

Panels

0

0

0

)

boom

access

Wing Access Panels Fwd of control surface linkage

panel

_J_I_IDN TDLERANCES UNLESS QTHERVZSE SPECIFIED

Routing

.XX + ,01 ,XXX + ,OOi +

I/2 °

EMBRY-RII))I)LE AERONAUTICAL UNIVERSITY I)AYTONA BEACH FLFIRII)A

TITLE WING ACCESS PANEL BRAWING ND. F93-2A-107-07

LAYOUT SHEET 1 1

, I!11'

1 Panel Flat Wrapped from top of rear spar to bottom of front sp

1 Panel Lower Surface from front spar to re spar

MS20450DD-8 Rivet Spacing tail

I 2.0"

boom

inborad

of

/

X

1 Panel Upper Surf, 1 Panel Lower Surfl from rear spar to

Ill.OUT

NOTE:

1 Panel Flat Wrapped :ram top of front spar :o bottom of front spar

FI:".P_, _'_

All wing panels fabricated of 2024-T5 Aluminum• 0.052" skin thickness outboard of tail boom. 0.05" skin thickness inboard of tail boom. No control surfaces show.

Wing Press

, Ltfl'

•°+++

_+

•°+••

••t

*+

•

Formed Tip

•+

MS20426DD-5 Rivet Spacing 1. 25" outboard of tail boom 1 Panel Upper Surface 1 Panel Lower Surface from front spar to rear

1 Panel Flat Wrapped from top of rear spar to bottom of rear spar spar

:e e

ailing

edge

]]IHF..NSZI3N Trtl ERANCES UNLESS I'ITHERVISE SPECIFIED

EEC.LPP_6

EMBRY-RIBBLE AERONAUTICAL UNIVERSITY DAYTONA BEACH FLORIDA

SCAL_ ,XX + ,01 ,XXX + ,001 +__i/2 °

DRA_/N BY J. VIEIRA

TITLE WING SKIN

PANEL

LAYOUT

IP(]U_OUT FPJU,4E i

Windshield

I

defrost

Ii

Occupant

air

ven

Rudder/Steering

. m

Elevator Air

Venting

.,t.I#OLDOUT

Engine

cooling

FRA_'_¢

air

"_.

t

flow

//

Heating\Cooling Cabin

3OARD

ventilation

air

air

pump

flow

PROFILE

DIMENSION TOLERANCES UNLESS OTHERWISE SPECIFIED

.XX :i: .01 .XXX :1:.001

DAYTONA

B 110-15 _ TITLE CONTROL DRAWING NO---F95-2A-

BEACH

FLORIDA

B'_'-'__7 DRAWN VIEIRA & SYSTEMS 10_

LAYOUT

_ CEPEDA

I _!/

Access

Rudder/Steering Elevator Air

Ilil'

Venting

panels

J i

X

IDATE 10-1.3

JDWG BY VIEIRA&:CEPEDA f

Engine

cooling

! DRAWING F93-2A-

oir

NO. 105-07

flow

I SHEET 2//2

_-DOUT

FRAME _'-

Typico

i................. ......... ........

Fibergloss nose

I Illl,

colqe

Bottom pGnel skin thick (All others ore 0.025 in

rivet

spacing

is

1 in

ii

;ss

is

0.040

in

])INEN..RIDN TOLERANCES UNLESS DTHERt/ISE SPECIFIE9

,XX + ,01 ,XXX ± ,001 +

i/2

°

EMBRY-RIBBLE AERONAUTICAL UNIVERSITY BAYT[]NA BEACH FLBRIBA

I)RAWN BY A. CEPEDA TITLE DRAWING

SKIN PANEL LAYOUT NO, F93-2A103-07

SHEET

2024-

lull I

¢3

56

MS20430DD-12

2024-

42

488

MS20430DD-5

2024-

41

320

MS20430DD-2

2024-

40

680

MS20430DD-8

2024-

39

2090

MS20426DD-5

2024-

38

4

NAS

1307-15P

STANDAF

37

4

NAS

1 308-4P

STANDAF

36

8

NAS

1305-1P

STANDAF

35

4

AN

34

64

33

2

INBOARD

BOTTOM

32

4

INBOARD

TRAILING

EDGE

PANEL

2024-

51

2

INBOARD

LEADING

EDGE

PANEL

2024-

30

4

CENTER

29

2

28

2

27

2

PRESS

26

1

REAR

25

1

FRONT

24

4

REAR

23

4

FRONT

22

4

SPAR

SHR

WEB

0.016"

THICK

2024-

21

4

SPAR

SHR

WEB

0.032"

THICK

2024-

STANDAF

12-17-J

NAS

STANDAF

1304-¢P WING

WING SKIN

LEADING

PANEL

OUTBOARD

2024-

PANEL

EDGE WING SKIN WING SKIN

2024-

PANEL

PANEL

FORM WING TIPS CARRY THROUGH CARRY THROUGH

202420242024202420242024-

LUG

2024-

LUG

4

SPAR

SHR

WEB

0.05"

19

4-

SPAR

SHR

WEB

0.125"

18

4

BUCKLE

STIFFNER

(#12)

2024-T3

17

12

BUCKLE

STIFFNER

(#11)

2024-T3

16

28

BUCKLE

STIFFNER

(# 10)

2024-T3

15

44

BUCKLE

STIFFNER

(#9)

2024-T3

14-

2

REAR

SPAR

CAP

NAS-544-02

2024-T3

15

2

REAR

SPAR

CAP

NAS-544-52

2024-T3

12

2

REAR

SPAR

CAP

NAS-544-55

2024-T3

11

2

FRONT

SPAR

CAP

NAS-54-4-10

2024-T3

10

2

FRONT

SPAR

CAP

NAS-,...344-50

2024-T3

9

2

FRONT

SPAR

CAP

NAS-544-69

2024-T3

8

10

WING

RIB

72.0"

CHORD

2024-T3

7

2

WING

RIB

65.0"

CHORD

2024-T3

6

2

WING RIB

61.5"

CHORD

2024-T3

5

2

WING RIB

58.0"

CHORD

2024-T3

4

2

WING RIB

51.4-"

CHORD

2024-T3

5

2

WING RIB

44-.7"

CHORD

2024-T3

2

2

WING RIB

CHORD

2024-T3

1

2

WING RIB

CHORD

2024-T3

ITEM

THICK

2024-T3

2O

THICK

MAT'L

DESCRIPTION

OTY

2024-T3

OR

PART

II

UHL.ESSI]TI.IO_SE

SPECIFIED

.XX _+ ,Ol ,XXX ± .001 ±

1/2 °

UNIVERSITY EMERY-RIDDLE AERONAUTICAL FLORIDA DAYTONA BEACH

SIZE B TITLE

DATE 12-06

SCALE INDICATED

DRAWN BY ALPHA TEAM

WING STRUCTURAL DRAWING NO, F93-2A104-07

LAYOUT SHEET

1/13

#

J_LI;)OUT

__o_,

fL

Wing 20

4-0

60

S

80

\

\

2_

L

Tail

Illll

B

tion O0

l,lll

(in.) 120

140

iI,lIi

160

180

2OO

,lll,l,l,l

Aileron

Fuel Tank

SCALE:

)m

ZDATE 10-15

IDW6 :BY TEAM ALPHA

1/20

I_RAWZNG F95-2A- Nn, 104-07

iSHEET 2/15

IPgu_UT F_

Vertical

MAIN

Tail

GEAR

SCALE:

' Illll

1/2C

.

Detail

1

Detail

2

Tail FUEL

T

IDATE 12-06

ID_/G BY ALPHA

TEAM

II)RAWING F95-2A-

NFI, 104-07

Boom

D

2.60

Ligh

l h

J "Detail

5

b Rib

DETAIL SCALE:

:3 1/2 Wing .-,--

IIIII

I

Skin 5.70

Stringe' in.

Rib

Location (in.)

ning

tched

and Chord

Dimensions (in.)

h

for

(in.)

NLF(1 )-041 b

(in.)

4

D (in.)

Hole

Rib

21.0

72.0

10.08

7.7

5.25

27.0

72.0

10.08

7.7

5.25

35.0

72.0

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