NASA-CR-I95488

_SU, u q_g_

/,,u'-o3-E,4J-

P- 7

COCKPIT CONTROL SYSTEM

=

421S93ADP01-2

03-12-93

AE421/03/DELTA

Lead Engineer: David Lesnewski

Russ M. Snow

Lisa Combs

Dave Paufler

George Schnieder

Roxanne Athousake

Submitted to:

Dr. J. G. Ladesic

(NASA-CR-195488) COCKPIT CONTROLSYSTEM (£mbry-Ridd]e Aeronautica]Univ.) 67 p

G3/08

N94-24957

Unclas

0204243

Q U

TABLE OF CONTENTS

ii°°°o ..........

pro|ect summary .............................................................................. ""

iii

1

Rudder

3

Design Description ......................................................................... 3

Loads and Loading ........................................................................

structural Substantiation .................................................................

Manutacturing and Maintenance Provisions ...................................

6

9

10

Eievat°rDesig n Description .......................................................................... 10

Adiustable control Y°ke .................................................................. 1_4

control Column Support ..................................................................

push-pull Rod Assembly .................................................................. 15

Elevator Trim .....................................................................................

Manufacturing and Maintenance Provisions ..................................... 16

17

Aiier°nSDesig n Description........................................................................... 17

Loads and Loading ......................................................... :..................

Structural Substantiation ................................................................... 2022

Manutacturing and Maintenance Provisions ..................................... 23

weight summary ......................................................................................

conclusiOnS .............................................................................................. 23

Appendix

List of Figures

Rudder

2.2.1 Toe Pedal ...................................................................... 3

2.2.2 Pedal Assembly ............................................................. 4

2.2.3 Pedal Mount ................................................................... 4

2.2.4 Pedal Actuated Crank .................................................... 5

2.2.5 Center Crank .................................................................. 5

Elevator

3.2.1 Adjustable Yoke ............................................................ 11

3.3.1 Control Column Support ............................................... 12

3.4.1 Push/Pull rods ............................................................... 15

Ailerons

4.2.1 Both Pilots Turning Inward .......................................... 17

4.2.2 Lower Pulleys ............................................................... 18

4.2.3 Both Pilots Turning Outward ........................................... 19

List of Tables

Stress Analysis

11.0 Rudder ...............................................................................

12.0 Elevator ..............................................................................

23.0 Ailerons ..............................................................................

Rudder

2.3.1 Toe Brake Pedal ............................................................... 6

2.3.2 Lower Pedal ...................................................................... 6

2.3.3 Pedal Mount ...................................................................... 6

2.3.4 Rod End Casting ............................................................... 6

2.3.5 Rod .................................................................................... 7

2.3.6 Rod End ............................................................................ 7

2.3.7 Pedal Actuated Crank ....................................................... 7

2.3.8 Bottom Mount .................................................................... 7

2.3.9 Top Mount ......................................................................... 8

2.3.10 Center Rod ...................................................................... 8

2.3.11 Center Crank ................................................................... 8

2.3.12 Center Crank Mount ........................................................ 8

Elevator

3.2.1 Pin Attachment Stress Summary ...................................... 11

3.3.2 Control Column Support Calculations .............................. 13

iii

1.1 PROJECT SUMMARY

The purpose of this project is to provide a detail design for the cockpit control system of the Viper

PFT. The statement of work for this project requires provisions for control of the ailerons, elevator, rudder,

and elevator trim. The system should provide adjustment for pilot stature, rigging and maintenance. MIL-

STD-1472 is used as a model for human factors criterion. The system is designed to the pilot limit loading

outlined in FAR part 23.397. The general philosophy behind this design is to provide a simple, reliable

control system; which will withstand the daily abuse that is experienced in the training environment without

excessive cost or weight penalties.

RUDDER

COMPONENT

Toe Pedal

Rod & Casting

Pedal Actuated Crank

fm_ (psi)

7541

4655

16076

MARGIN OF SAFETY

1.65

3.30

.7

Bottom Mount 6024 2.32

Center Cast 8639 2.16

17278 .15Center Mount

PAGE #

Appendix

Appendix

Appendix

Appendix

Appendix

Appendix

Table 1 : Rudder Analysis

ELEVATOR STRESS ANALYSIS

FITTING FACTOR MARGIN PAGE #

fmx FACTOR fL" OF fUL OF(psi) SAFETY (psi) SAFETY

Tw

Support 15504 1.2 18605 1.5 27908 1 12Column

902 1.2 1082 1.5 1.51 14Bolt Case1

BoltCase 2

Joiner

2204

1803 1.4

4404

2524

1.5

2

1623.6

6606

5048

.41

10.1(Y)6.13(U)

"" I ne margin ot satety at yield was only calcuiated tor the weakest spot m, t_e joiner, see drawing AI=421:593AUP01-3

14

14

ControlYoke(s)

ControlYoke(brg)

Rods

f.= FITTING(psi) FACTOR

4074 2

13250 2

11459 1.2

iLL

(psi) OFSAFETY

8146 1.5

26500 1.5

13571 1.5

(psi)

12222

51200

27501

Table 2: Elevator Stress Analysis

MARGINOFSAFETY

4.64

.4

1.04

PAGE #

11

11

15

COMPONENT

Sprocket 1 & 2

Sprocket 3

Pulleys 5 & 6

AILERONS

f=x (psi)

21281

3396

17195

MARGIN OF SAFETY

1.97

17.74

PAGE #

20

21

3.70 21

Table 3: Aileron Stress Analysis

2.0 RUDDER CONTROL SYSTEM

2.1 Description

The rudder control system provides rudder and nosewheel deflection of +-20° as well as differential

braking. The differential braking is composed of a separate master cylinder, parking brake and valve (see

figure in appendix). Pilot input is transferred from two sets of bottom pivoting rudder pedals, through a

series of push-pull rods and bellcranks, which are routed through the driveshaft tunnel.

2.2 Loads and Loading

2.2.1 The toe brake structure pivots on a bearing at point

A. A 200 lb. load supplied by the pilot generates reactant forces

at A and B. The following calculation is when the master cylinder

is fully compressed.

Summing the moments about point A;

.__b s.

Figure 2.2.1

_- 0

2001bs. x 2.00in. - FB x 1.75in. - 0F_ u

228.61b

Summing the forces in the X and Y plane;

Fy- 0

F _vr ,2+ ,,

FAx- 2001bs. - 228.61bs. x sin9.4 - 0

228.61bs. x cos9.4 - FAy " 0

FA- 282.81bs.

2.2.2 The largest value for Fc and F D occur when the

pedal is in the neutral position. A pilot's force of 200 Ibs. is

applied at the toe brake to create the largest moment.

Summing moments about point C and the forces in the

horizontal plane;

FD x 2.75in. - 2001bs. x7.75in. - 0

200 Ibs.

r

Figure 2.2.2

FD - 563.6 ibs.

F c - 2001bs. - 563.61bs. - 0

F c- 763.61bs.

2.2.3 Pedal Mount

- 4Fbolt - 0 Fbolt - 1411bs.

Figure 2.2.3

2.2.4 Rod Assembly

A load of 763.4 Ibs. is carried through the rod connecting casting, rod, and the rod ends that attach

to the pedal actuated crank.

2.2.5 Pedal Actuated Crank

The bell-cranks' maximum load occurs when the pilot

uses full braking on both pedals; therefore, FCENTERand FGEAnare

equal to 0.

763.6 Ibs . + 763.6 ibs .-FpzvoT - 0

---,--C:3Z_

Figure 2.2.4

Fp_voT - 1527 ibs.

2.2.6

pedal.

Center Rod

The maximum axial load is 736.6 Ibs.. This occurs with complete resistance to pilot load on one

2.2.7 Center Crank

The limiting case occurs when both pilots act in unison.

_ M_,zvo_, - 0

FR_ - 15271bs.

Fcm_r_2 .5 - Fao_2 .5 - 0

FpIvoT" _15272 + 15272 Fpzvo T " 21601bs.

2.5_m.

® --J

Figure 2.2.5

5

2.3 Substantiation

Toe Brake Pedal

Figure 2.2.1

Point A

Point B

t (in)

.30

.25

F.F. f_.,

2.0 7541

2.0 7315

Table 2.3.1

f to. M.S. LL

3771 1.65

3658 1.73

I M.S.u.L

1.83

1.92

Toe Brake Pedal

Lower Pedal

Figure 2.2.2 t (In) d (In)

Point A

Point B

Point C

Point D

.25 .656

.25 N/A

1.79 N/A

.25 .75

fb,_ f to. fo M.S. LL(psi) (psi) (psi)

1724 N/A N/A 10.6

3658 1829 N/A 4.47

3413 1706 N/A 4.86

N/A N/A 638 30.4

M.S.U.L

11.4

4.83

5.25

32.4

Table 2.3.2 Lower Pedal

Figure 2.2.3

t(in)

1.0

Pedal Mount

d(In) f_ fro.(psi) (psi)

I1.3 I 434

I

Table 2.3.3 Pedal Mount

1127

M.S,LL

16.7

M.S.u. L

17.9

t(in)

.25

d(in)

.656

Rod End Casting

fbqll

(psi)

i

M.S.LL M.S.u. L

4655 3.3 3.94

Table 2.3.4 Rod End Casting

,i

Length (In) A(In "=) I(In'4)

6.00 .06943 .001786

Rod

fcomp f©rlt M.S.LL

(psi) (psi)

10998 2116 5.82

Table 2.3.5 Rod

M,S,u. L

4.76

I Rod End

FLL (Ibs) I M'S'LL M.S.u.L

3025 I 2.96 1.64 li

Table 2.3.6 Rod End

Figure 2.2.4

Pedal Actuated Crank

t (In) fb,Q(psi)

•19 16076

f tO.

(psi)M,S, LL M.S.u. L

Point A & C 4257 1.55 .70

Point B .19 6.65 4.1

Point D .19 2.64 1.43

Point E .19 1.55 .70

5358 2297

11253 2993

16O76 4019

Crank Assembly In Bending

Moment C I S.C fa,,_,o M'S'LL M.S.u.LArm(in) (In) (In") Factor (psi)

7.0 1.25 .261 3.0 I 40512 1.23 .053

Table 2.3.7 Pedal Actuated Crank

7

BottomMount

ft_. f_ f to. M.S. LL M'S'u.L

(psi) (psi) (psi)

Pivot 4122 6024 2447 2.23 2.54

Bolts N/A 3054 3054 5.55 5.98

Table 2.3.8 Bottom Mount

Top Mount

ft,,, f_(psi) (psi)

4122 6024

M.S. LL M'S'u.L

2.23 2.54

Table 2.3.9 Top Mount

Center Rod

Area Length(In `=) (In)

.06943 9.00

I f¢omp

(In`4) (psi)

.001786 10998 I

for. M'S'LL M.S-u.L(psi)

94000 5.82 4.76

Table 2.3.10 Center Rod

i

Figure 2.2.5

Point A & C

Point B

ften,,

(psi)

N/A

5759

Center Crank

f_(psi)

24435

8639

f t.O.

(psi)

8551

5759ii

Table 2.3.11

M.S, LL

.678

3.75

M,S.u. L

.119

2.16

Center Crank

I CenterCrankMount

ftl.= fbro f to.

(psi) (psi) (psi)

Pivot I 2160 17278

M.S. LL M.S-u.L.

1439 .158 .235i

Table 2.3.12 Center Crank Mount

2.4 Manufacturing and Maintenance

Aluminum castings are utilized in the manufacture of most components in the system. The casting

alloy is 355.0-T6, which is a light weight, high strength casting alloy. The pedal actuated crank and center

crank are fabricated from 2024-'1"6aluminum plate. The push pull rods are one half inch O.D. 4130 steel

tubing. All other components are aircraft approved vendor supplied hardware, includingthe braking system.

3.0 ELEVATOR CONTROL SYSTEM (ECS)

3.1 Description

Theelevator control system was designed using the loading conditions listed in FAR Part 23 and

the human factors specifications given in MIL-STD-1472 Appendix Two. The first step in the design

analysis was to determine how the elevator control system would function. As is typical in most aircraft,

the pilot would exert a force on the control yoke, which would transfer the motion horizontally to the control

support column, which would then transmit the load to push pull rods, which connect to a bellcrank with

dual universal joints that were connected to yet another rod which facilitates the proper deflection angle

with a maximum angle of ___20°. The next step was to organize the elevator control system into workable

sections. These sections are as follows:

1. The adjustable control yoke shaft2. The control column support (including the pivot point)3. The actual push-pull rod assembly with interface to the elevator4. The elevator trim

After defining each section of the ECS, the sections were thoroughly researched using typical single

engine models like the Cessna 172, the Mooney and the Piper Cadet(Warrior). With this accomplished,

it was time to begin the structural substantiation and material selection. For clarity, each of the sections

mentioned above will be discussed individually.

3.2 Adjustable Control Yoke

The control yoke with adjustable arm was a unique idea conceived for the discriminating pilot who

is not of "standard" height and arm length. The adjustable control yoke shown below, allows the pilot to

achieve proper elevator deflection by compensating for his/her arm length. The control yoke can be made

longer for a pilot with long arms and shorter for a pilot who has shorter arms.

This increases the pilot's comfort as well as his safety, because proper deflection can be

made. The control yoke is adjusted by pushing in and simultaneously turning the fastener and moving the

10

inner shaft to the proper location, which is

determined by the pilot's arm length, then

releasing the fastener, which allows the steel pin

to lock into position. This adjustment method was

chosen primarily for its safety features. With the

Figure 3.2.1

two motions required to relocate the pin, it seems unlikely that pin will be able to displace if bumped or

knocked during flight.

3.2.1 Substantiation

The adjustable control yoke was designed using the worst case scenarios of loading, ie. a 2001b

force is exerted by the pilot, which is represented in figure 3.2.1. The 200 Ib load will be transmitted

directly through the control yoke shaft to the control column support. Of major interest is the pin attachment.

The table below summarizes the results of the calculations.

Pin Attachment

fSheer

(psi)

4074

f_ F.F. F.S. M.S.LL M.S.u.L(psi)

12800 2.0 1.5 9.06 I .40

Table 3.2.1 Pin Attachment Stress Summary

11

3.2.2 Manufacturing and Maintenance

The inner and outer shafts are made of extruded Aluminum 2219 and are engaged

such that buckling will not be a problem. Located at the intersection of the dash and the control yoke is

casted "dash interface" designed to provide horizontal support. A universal joint is located approximately

6 in from the interface to allow the control yoke to move horizontally with no vertical displacement, while

the control column support moves radially 10.9 ° . The displacement of the control yoke is negligible as

shown from the following calculations.

(5= fl/G = 200*33/3843197 -.) (5= .0017 in.

The control yoke is welded to the control column support with a needle-thrust point ball bearing to facilitate

turning motion necessary for the ailerons.

3.3 CONTROL COLUMN SUPPORT

Essential to the ECS is the control column support shaft, shown in

figure 3.3.1. The control column support shaft function is to pivot around the

base point when the control yoke is pulled out or pushed in, thus transferring

motion to the push-pull rods located approximately 3 in. above the base point.

r i3

D

3.3.1 Loads and Loading r •

Figure 3.3.1This part was designed for the two worst case scenarios that the

support column will encounter; 1) both pilots pushing or pulling with a 200 Ib

force each or 2) one pilot pulling while the others pushes with a 200 Ib force each. The support column

was analyzed by breaking it into three members, 2 vertical and 1 horizontal. The static analysis is shown

12

below,where H = Reaction load and E = load that the push-pull rod will see.

Z;Fx= 0 2*200 - E - H = 0 T_My= 2"200"16.8 - E - H = 0 --> E = 2203 Ibs.

The results for the members are tabulated below.

H = 1803 Ibs.

3.3.2 Substantiation

Case Member Shear(psi) Moment Torque

#1 200 1980

# 1 #2 400 4520 1980

#3 "2203 "5179,.. ,n

#1 200 1980

#2#2 200 2260 3960

0 0 "4520toL)Orlotes _lfllUftl va

#3_S that were design(

Table 3.3.2 Control Column Support Stress Analysis

The calculations shown below substantiate that the control column support can made of extruded Aluminum

2219 and that AN4C-24 bolts can be used to connect the control column joiner to the control column

attachment plate.

fn_l = TC/J= 4520"1/.57

f.=l = 2879 psi

f.=2 = MC/I= 5179"1/.785

f_== = 6597 psi

%=,3= VQ/It= 2203".537/.785".25

f.=_ = 6028 psi

T.F = f_axl + frn,x2+ f,=3 = 15504 psi --) fLL= 1.2*Z;fr_ --) M.S._ = 36000/18605 - 1 = .93

") fuL = 1.5*fLL ") M.S.=, = 56000/27908 - 1 = 1.0

13

The maximum load for the bolts will be found at point E ( push-pull rod location). Four bolts will

have to sustain 3606 psi. Thus, each bolt will see 902 psi.

T Fx = 0 V -1803/2 = 902 psi -) fCL= 1.2"V = 1082 psi -> fuc= 1.5*fLL = 1623.6 psi

M.S. = (Ftu / fu_- 1 = 4080/1623.6 -1 = 1.51

The weakest spot will be where the control column attachment plate, which will be

machined as well as the control column joiner, is welded to the aircraft structure. The

substantiation calculations are shown below.

-.)

f=P/A F.F.= 1.2+.2(weld) fcL=F'F*f= 1803/2".5 = 1.4 = 1.4"1803

f = 1803 psi F.F. = 1.4 fLc= 2524 psi

fUL= F.S. * fLL= 2.0 * 2524

fUL= 5048 psi

M.S.y,,td = 36000/5048 - 1M.S._,_ = 6.13

M.S._I = 56000/5048 - 1M.S.., = 10.1

3.4 PUSH-PULL ROD ASSEMBLY

The push-pull rods were selected for the D-2 for several reasons. First the push-pull rods have

a "better" feel to them thus the student pilot has a better idea of the deflections necessary to create proper

elevator motion. Second, the push-pull rods, have a greater coefficient of thermal expansion than the

traditional cable. This is an important factor considering that the ECS is located several inches below the

engine. These push-pull rods are connected to the control column support by means of a NAS-660-R4-11

rod end. The rods are extruded 4130 Steel and will see a load of 2200 Ibs, see control column

calculations. The rods will be supported approximately every 48 inches, using a fabricated idler arm

assembly(see rudder). The calculations for this are shown directly below.

14

3.4.1 StructuralSubstantiation

f = P/A fu = 1.2"f= 2200/.616 = 1.2"11459

f = 11459 psi f,, = 13751 psi

fuL = F.S. * fLL= 2 * 13751

fUL= 27501 psi

M.S.y,,id = Ft_Jfty-1= 39000/13751-1

M.S_,_ = 1.84

M.S.., = FtJft.-1= 56/27.5 -1

M.S. u,= 1.04

Pc_ = _^2 ,E,I/I.,2 = _,_,(30E6),(.001766)/L^2 _) L= 48.6 in

The push-pull rods transfer the

displacement of the control column support to a

bellcrank (see rudder) has dual universal joints on

f--22001bs_ _1 f--22001blJ

Figure 3.4.1

MS20271-B10 on either end which oscillate the elevator around two internal hinges. These U-Joints were

selected according to the maximum torque that was calculated for the pivot point.

3.5 Elevator Trim

The elevator trim was conceptualized as a cable and pulley system attached to a jackscrew that

controls the elevator trim deflection. The system is lightweight and is not used as frequently as the

elevator, therefore the "feel" is not as important and using push-pull rods cannot be justified. According

to the MIL Handbook, a pilot can only exert 59 Ibs by gripping the trim wheel and turning it. It was then

decided that in the interest of producibility, the trim system would use the same pulleys, sprockets and

cable as the aileron system, which was designed for 160 Ibs. This creates a factor of safety of

approximately 4.0 for the elevator trim system. The trim wheel and indicator are located on the control

panel in between the pilot's seats and can be rotated clockwise or counterclockwise which turns the cable

and, producing motion in the jackscrew, which connects to a piano hinge to create a + 10° maximum

deflection.

15

3.6 Manufacturing and Maintenance

The entire elevator control system was designed with maintainability and life cycle in mind. Access

panels are located on the bottom portion of the fuselage and on the cowling to provide easy access to both

the push-pull rods, the trim system and the control column support.. The stretch in the cable can be

adjusted using the turnbuckles and cable joiners. The motion of the part and its use were kept in mind

while selecting the proper material to withstand the stresses. The same part was used throughout the

cockpit control system, ie the idler arms were used for both the rudder and elevator, to minimize the

number of parts that have to be stocked or ordered. In conclusion, the elevator control system meets the

requirements of the SOW for the Viper D-2, FAR Part 23 and MIL Handbook. The pin attachment was

originally conceptualized as a cog wheel track with a locking pin but was changed to the pin idea

presented in this report.

16

4.0 AILERON CONTROL SYSTEM

4,1 Description

The aileron control system provides aileron deflections of positive 20 degrees an negative 15 degrees.

The system is operated by rotating the control yoke 90 degrees in either direction; clockwise for right roll

and counter-clockwise for left. The cockpit portion of the system using sprockets and chains to transfer

yoke rotation to the lower portion of the control column. This is translated through cables which run

through the driveshaft tunnel to a set of bellcranks which transfer displacement to the ailerons.

4.2 LOADS AND LOADING

The analysis of the "cockpit "portion of the aileron system will begin with the determination of the

static loads created by the application of the forces which correspond to the maximum possible forces as

prescribed by the federal aviation administration.(F.A.R 23.397) It will also be taken into consideration that

the system has been preloaded to 115 percent of these F.A.R values prior to the current application of the

forces. It should also be noted that these forces will be added to the system in a way as to simulate the

worst possible condition that could occur in the flight environment.

17

4.2.1 The first of the two worst case scenario's includes both pilots turning their wheels inboard

simultaneously at the prescribed torque. The resultant static loads are shown below. Remember each

closed system is preloaded to 620 pounds.

Summing the forces at sprocket $1 to find the

resultant force on the bolt.

T1 =T2=81 Ib

T3=T4= 11591b

T5=T6=6201b

_Fysl--COS (48.66) 81-C0S (48.66) II.

• 5391bs i II 442 IN

34/£ l

Figure 4.2.1

S1

51;39 Ibs

Fal,sl-819.05 ib

_'Fxsl--SIN(48.66 )81-SIN(48.66 )1159 +Faxsl

Faxsl " 9 31. O lbs .

F_-124HxOOlb

By symmetry the resultant force on sprocket $2 is equal to 81

Fas2 - 1240.01bs.

18

Summing the forces at sprocket $3 to find the resultant force on the bolt.

• F s- (819) 2- (57 0) 2+Fa_.s3

FRysj- FRs3-49 8lb

Summing the forces, in the plane parallel to the

pulley, at pulley P5 to find the resultant force on the

bolt.

_'Fx'COSII. 83 (620) +Faxp5

FR_s-606. 831b

_Fs-620-SINII .83 (620) -Fan_5

= P1

14.75 IN

Figure 4.2.2

Fa_s'-492.891b

F_5-7 81.78

19

4.2.2 For the second worst case scenario

the yoke wheels are both turned outward

simultaneously with the same torque the

values. In this case T1,T2,T3, and 1"4 will

swap values but will result in identical resultant

values on the bolts.

. . 539 lbs i_ 11.442 IN. S/15_

S2 '- _-" 59 lbs

Figure 4.2.3 i

4.3 STRUCTURAL SUBSTANTIATION

Now that the system has been statically defined these values will be used to evaluate the structural

fortitude of the members that are recipients of these loads. This will be done by evaluating the various

stresses and the corresponding factor and margin of safety. It should be noted that all forces have been

multiplied by a fitting factor of 1.5.

20

4.3.1 Sprockets 1 and2.

fshear" P 1860 16909 09 ibs-- w 1,m •

A 0. II in.2

fbz " _ " 1860 - 21281.46 ibs.td 0.23 (0.38) in.2

Due to bearing stress:

M.S.L.L. - 1.97, F.S.r,.L" - 2.97, M.S._.L. - 3.45, F.S.u.L. - 4.45

4.3.2 Sprocket assembly 3.

P 498.1 " 3396.00 Ibs.fsh6ar " _ " 0.22 in.2

P 498.1 - 585.17 ibs-----ufb,aring" t-d" 1.60(0.53) in.2

Due to shear:

M.S.r,.r..- 17.74, F.S.r..L" - 18.74, M.S.u.r." - 27.11, F.S.u.L. - 28.11

21

4.3.3 Pulleys 5 and 6

P 1172.67 , 13029.67 ibs.fsh_,_ " _ " 0.09 in.2

P 1172.67 - 17194.57 ibs.fbearlng " t-d " O. 20 (. 341) in. 2

Due to bearing:

M.S.T..r. " 3.70, F. SL.L. - 2.70, M.S.u.L. - 4.55, F.S._.L. - 5.55

4.4 Manufacturing and Maintenance

The aileron control system consists largely of aircraft approved, vendor supplied hardware; including

all fasteners, fittings, chain, and cable. Manufactured components include the bellcranks and female jack-

screw assembly which are cast from 355.0-T6 aluminum. The sprocket assemblies require welding which

will be done in-house.

22

5.0 WEIGHT SUMMARY

I!COMPONENT

I[ WEIGHT (Ibs)

Rudder

Pedals 4.1

Bellcranks 2.0

Brakes 3.0

Rods 4.4

Hardware 7.1

Elevator J

Adjustable Control Yoke Assembly 2.9

Control Column Support Assembly 3.1

Push-Pull Rods 4.2

Trim 3.2

Ailerons

Chains 3.1

Sprockets .49

Pulleys .9

Cable 6.0

Bellcrank 4.0

TOTAL Jl 47.8

6.0 CONCLUSION

The goal of this design project was to provide a simple yet effective cockpit control system, meeting

FAR specifications and designed with human comfort in mind. It has been shown that his design meets

the above listed criteria, but their is always room for improvement. Several components have been

designed with excessive margins of safety, and weight reduction is possible by removing excess

material.

23

APPENDIX

Design substantiation was done for all cockpit control system components. Analysis consisted of

Margin of Safety calculations for the limit load and ultimate loading cases. Fitting factors were applied to

both rotating and non-rotating joints. It should also be noted that a stress concentration factor of :3.0 was

applied to the pedal actuated crank in bending. The following equations were used in the substation

analysis.

P--E_arlng " t---dx F. F

Pft.o. " X F.F.

2Xt

f _ens " P x F. F.(h-d) t

P

fshear " _ x F.F.

fczlc.._.l_2EI

A12

F. F. ro_atlng " 2.0 F. F. non-rotatlng " 1.2

1eld FULT 1M.S.r..L. " fr..L. 1 M.S. ozT" fL.L.I.5

BRAKE SYSTEM

PILOT'S MASTER

CYLINDERS

HYDRAULIC FLUID

RESERVOIR

COPILOT'S

MASTER CYLINDERS

\

PARKING

BR.AKE VALVE

PARKING BRAKE CONTROL

L.H BRAKEl_H BRAKE

u

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22 i

21 26

20 12

19 4

18 2

17 2

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BDLT

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16 2

15 2

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18 8

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10 2

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421S9303DI8E

42139303DI8E

TERMINAL

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BOLT

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3/16 FLEXIBLE 3]'EELCABLE

CABLE, CHAIN UNION

MIL-C-5424

42139303D180

07 1 TRIPLE SPROCKET ASSEMBLY 421S93031)17E

061 BEARING .... 421S9303I)176

i05 1 ,5321N DIA, 4,8831N LEN, STEEL 421S93031]174

04 2 SPRnCKET SUPPORT BRACKET 421S9303D172

03 5 0,500 PITCH SPROCKET 41EM-B-12

102 2 ,,3791N BIA, 1,OOIN LEN, STEEL 421S9303D170

I011 601N 0,500 PITCH CHAIN RC4155 ....I

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0713032-1

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IDLER ARM ASSEMBLY 421S93D10306

_USH PULL RODS.; 4130 STEEL .5=D, .04=T

PUSH PULL ROD ENDS NAS-660R411

CONTROL COLUMN ATTACHMENT PLATE ALUM 2219

CONTROL COLUMN JOINER ALUM 2219

CONTROL COLUMN SUPPRT SHAFT ALUM 2219

CONTROL COLUMN U-BAR AMS-4162

NEEDLE-THRUST POINT BALL BEARING NKX30/DUST

UNIVERSAL JOINT MS20271-310

COLLAR 0713032-2

YOLK SHAFT FASTENER 5610-401-20B

ADJ, YOLK SHAFT ALUM 2219

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YDKE PIN05-16610

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36

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32

31

30

29

28

27

26

25

24

23

22

21

20

19

18

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BEARING

IDLER ARM HDUNT

BALL JOINT

CENTER CRANK HDUNT

ROD END CASTING

LDWER PEDAL

UPPER PEDAL" ,.. m.

RDD END

VALV

01

03

02

O1

04

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04

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19

RESERVOIR

20 BRAKE LINE PER FDDT

O1 SCOTT PARKING BRAKE

Ol BRAKE

04 CLEVELAND HASTER CYL,

08 BEARING INNER RACE

08 BEARING

22 BEARING

75 AIRCRAFT WASHER

75 CASTLE NUT

19 AIRCRAFT BOLT

04 AIRCRAFT BDLT

32 AIRCRAFT BDLT

IDW

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AN2

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18 AIRCRAFT BOLT AN4-12

02 AIRCRAFT BOLT AN4-16

02 CIR, ROD 4130 9 IN, 421S9303D14,

06 CIR, ROD 4130 6 IN, 421S9303D12_

Ol GEAR BELL CRANK 421S9303D12_

O1 TORQUE TUBE ASSEMBLY 42139303D12.

Ol CIR, ROD 4130 23 IN, 421S9303D12_

Ol REAR BELL CRANK 421S9303D12[

Ol CIR, ROD 4130 50 IN, 421S9303DI18

Ol CIR, ROD 4130 37 IN, ,421S9303DIIE

03 IDLER ARM 421S9303DI14

O2

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CIR, ROD 4130 51 IN;II

RGHT PEDAL ACTUAT, CRANK

LEFT PEDAL ACTUAT, CRANK

CENTER CRANK ASSEMBLY

421S9303DI12

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421S9303D108

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