J

.

I,JSRA

HIGH ALTITUDE

RESEARCH AIRCRAFT

VOLUME I

SUBMITTED - 1 JULY 1990
08WF INDU8TRIE8

CALIFORNIA STATE

POLYTECHNIC UNIVERSITY

POMONA

NASA/USRA

HIGH ALTITUDE RECONNAISSANCE AIRCRAFT

GROUP MEMBERS

MICHAEL RICHARDSON

JUAN GUDINO

KENNY CHEN

TAI LUONG

DAVE WILKERSON

ANOOSH KEYVRNI

MICHAEL MEDICI - PROJECT ENGINEER
TABLE OF CONTENTS

1.0

2.0

3.0

4.0

5.0

6.0

INTRODUCTION

!.! REQUIREMENTS AND CONSTRAINTS

1.2 MISSION PROFILE

INITIAL DESIGN

2.1 DESIGN CRITERIA

2.2 CONFIGURATIONS

FINAL DESIGN

3 I PRELIMINARY WEIGHT ESTIMATION

3 2 WING GEOMETRY

3 3 HORIZONTAL TAIL

3 4 VERTICAL TAIL

3 5 FUSELAGE

3 6 REFINED WEIGHT ESTIMATION

3 7 MOMENTS OF INERTIA

3 8 CENTER OF GRAVITY LOCATION

AERODYNAMICS

4.1 AIRFOIL SELECTION

4.2 LIFT

4.3 DRAG

STABILITY AND CONTROL

5.1 STATIC STABILITY

5.2 DYNAMIC STABILITY

5.3 FLYING QUALITIES

PROPULSION SYSTEM

6.1 SYSTEM REQUIREMENTS
7.0

8.0

6.2 POWERPLANTSELECTION

6 2 1 TURBOJETS/TURBOFANS

6 2 2 TURBOPROPS

6 2 3 HYDRAZINEENGINE

6 2 4 INTERNAL COMBUSTIONENGINES

6 2 5 OTHERTYPES OF POWERPLANTS

6 2 6 SELECTION OF THE POWERPLANT

6.3 ENGINE CONFIGURATION

6 3 1 TURBOCHARGING SYSTEM

6 3 2 ENGINE BLOCK AND CYLINDERS

6 3 3 LUBRICATION SYSTEM

6 3 4 COOLING SYSTEM

6 3 5 FUEL SYSTEM

6 3 6 IGNITION SYSTEM

6 3 7 GEAR REDUCTION

6 3 8 ELECTRICAL SYSTEM

6 3 9 ENGINE CONTROL SYSTEMS

6.4 PERFORMANCE SPECIFICATIONS

6.5 PROPELLER DESIGN

PERFORMANCE

7 1 TAKEOFF

7 2 LANDING

7 3 CLIMB

7 4 CRUISE

7 5 FLIGHT ENVELOPE

? 6 POWER REQUIRED

STRUCTURES
9,0

I0 0

II 0

12 0

13 0

14 0

15 0

16 0

!7 0

8 !

8 2

8 3

8 4

8 5

8 6

8 7

8 8

STRUCTURAL ANALYSIS

FINITE ELEMENT ANALYSIS

CONSTRAINTS

LOADING OF THE MODEL

DEFLECTION

TSAI-HILL FAILURE CRITERIA

STRESS

BENDING MOMENT

LANDING GEAR

9 1 LANDING GEAR ARRANGEMENT

9 2 RUNWAY LOADS

9 3 LANDING GEAR TIRES

9 4 SHOCK ABSORBERS

9 5 GEAR ST_VUTURE

9 6 LANDING GEAR HYDRAULIC SYSTEM

9 7 BRAKE SYSTEM

RELIABILITY AND MAINTAINABILITY

COCKPIT VISION AND HUMAN FACTORS

COST ANALYSIS

MANUFACTURABILITY

ELECTRICAL SYSTEM

LIFE SUPPORT

CONCLUSION

RECOMMENDATIONS

REFERENCES

APPENDIX A

REQUEST FOR PROPOSAL
APPENDIX B

COST BREAKDOWN

APPENDIX C

MAINTAINABILITY & AVAILABILITY ANALYSIS
Symbol

b

Cd

c.g.

c!

CL

Clmax

Cla

CLa

Cmc/4

C-F

C-D-alpha

C-D-u

C-L-i-h

C-M-i-H

C-L-alpha

C-L-u

C-L-q

C-L-delta-E

C-L-alpha-dot

LIST OF SYMBOLS

Definition

Wing Span

Total Aircraft Drag Coefficient

Center of Gravity Location

Airfoil Lift Coefficient

Total Aircraft Lift Coefficient

Maximum Airfoil Lift Coefficient

Airfoil Lift Curve Slope

Wing Lift Curve Slope

Pitching Moment Coefficient About

Airfoil Quarter Chord

Turbulent Skin Friction Coefficient

Variation of Drag Coefficient with

Angle of Attack

Variation of Drag Coefficient with

Speed

Variation of Lift Coefficient with

Stabilizer Incidence

Variation of Moment Coefficient

with Stabilizer Incidence

Variation of Lift Coefficient with

Angle of Attack

Variation of Lift Coefficient with

Speed

Variation of Lift Coefficient with

Pitch Rate

Variation of Lift Coefficient with

Elevator Deflection

Variation of Lift Coefficient with

Angle of Attack Rate
Symbol

C-m-alpha

Attack

Definition

Variation of Pitching Moment

Coefficient with Angle of

C-m-q Variation of Pitching Moment

Coefficient Pitch Rate

C-m-delta-E Variation of Pitching Moment

Coefficient with Elevator

Deflection

C-m-alpha-dot Variation of Pitching Moment

Coefficient with Angle of Attack

Rate

C-y-beta Variation of Side Force Coefficient

with Sideslip Angle

C-y-p Variation of Side Force Coefficient

with Roll Rate

C-y-r Variation of Side Force Coefficient

with Yaw Rate

C-y-delta-R Variation of Side Force Coefficient

with Rudder Deflection

C-l-beta

Angle

Variation of Rolling Moment

Coefficient with Sideslip

C-l-p Variation of Rolling Moment

Coefficient with Roll Rate

C-l-r Variation of Rolling Moment

Coefficient with Yaw Rate

C- l-delta-A Variation of Rolling Moment

Coefficient with Aileron Deflection

C-l-delta-R Variation of Rolling Moment

Coefficient with Rudder Deflection

Cpn-beta

Angle

Variation of Yawing Moment

Coefficient with Sideslip

C-n-p Variation of Yawing Moment

Coefficient with Roll Rate

C-n-r Variation of Yawing Moment

Coefficient with Yaw Rate
Symbol

C-n-delta-A

C-n-delta-R

CHAR. EQU.

FT

FT/S

hp

HORZ.

Ixx

Iyy

Izz

lbs

LBS

mm

M

MAC

OEI

Re

RN

S

SG

SR

STR

SCL

S A

SFR

Definition

Variation of Yawinq Moment

Coefficient with Aileron Deflection

Variation of Yawing Moment

Coefficient with Rudder Deflection

Characteristic Equation

Feet

Feet per Second

Horsepower

Horizontal

Rolling Moment of Inertia

Pitching Moment of Inertia

Yawing Moment of Inertia

Pounds

Pounds

millimeter

Mach Number

Mean Aerodynamic Chord

One Engine Inoperative

Reynolds Number

Reynolds Number

Planform Area

Ground Roll Distance

Roll Distance

Transition Distance

Climb Distance

Air Distance

Free Roll Distance
Symbol

SB

Sref

t/c

T

VERT.

Vsta!l

Vcrit

Vto

W

Definition

Braking Distance

Reference Area

Airfoil Thickness

Ratio

Thrust

Vertical

Stall Velocity

Critical Velocity

Take-off Velocity

Weight

to Chord Length
ABSTRACT

At the equator, the ozone layer ranges from approximately

80,000 to 130,000+ feet which is beyond the capabilities of the

ER-2, NASA's current high altitude reconnaissance aircraft. The

Universities Space Research Association, in cooperation with NASA,

is sponsoring an undergraduate program which is geared to designing

an aircraft that can study the ozone layer at the equator. This

aircraft must be able to cruise at 130,000 feet for six hours at

Mach 0.7 while carrying 3,000 ibs. of payload. In addition, the

aircraft must have a minimum of a 6,000 mile range. The low Mach

number, payload, and long cruising time are all constraints imposed

by the air sampling equipment. In consideration of the novel

nature of this project, a pilot must be able to take control in the

event of unforseen difficulties.

Three aircraft configurations have been determined to be the

most suitable for meeting the above requirements, a joined-wing, a

bi-plane, and a twin-boom conventional airplane. Although an

innovative approach which pushes the limits of existing technology

is inherent in the nature of this project, the techniques used have

been deemed reasonable within the limits of 1990 technology. The

performance of each configuration is analyzed to investigate the

feasibility of the project requirements. In the event that a

requirement can not be obtained within the given constraints,

recommendations for proposal modifications are given.
!.0 INTRODUCTION

In 1974, F. Sherwood Rowland and Mario Molina, chemists

at the University of California, theorized that the ozone

layer which protects the earth from harmful ultra-violet

radiation was being destroyed by chloroflourocarbons.

Chloroflourocarbons, or CFC's as they are commonly referred

to, are released into the atmosphere from sources like

refrigeration systems, styrofoam production facilities, and

aerosol cans to name a few.

Rowland and Molina's theory was met with great

skepticism by the scientific community when first published.

Scientists as well as the general public had difficulty

believing that the Earth's survival was being threatened by

the use of hair spray and hamburger containers. Now in the

1990's, the evidence accumulated over the last two decades

seems to support Rowland and Molina's theory, the Earth's

precious ozone layer is disappearing.

Because of the potential consequences of a depleted

ozone layer, scientist are desperately trying to investigate

this phenomenon. They are however, limited by the present

methods of collecting ozone data. Ninety percent of the

ozone layer lies 50,000 to 115,000 feet above the earth's

surface. NASA's highest flying atmospheric sampling

airplane is the ER-2 which has a service ceiling of only

70,000 feet. Clearly the ER-2 would not be able to sample

the majority of the ozone layer.
An alternative to the ER-2 is a large weather balloon.

Weather balloons are capable of reaching altitudes of

115,000 feet and beyond but they lack the directional

control required for sampling specific target areas. Still

another alternative would be rockets carrying sampling

equipment. Rockets, however, fly at Mach numbers that are

not compatible with current atmospheric sampling equipment.

Clearly there is a need for an aircraft that can

effectively sample this region of the Earth's atmosphere.

It is for this reason that NASA and the USRA developed a

request for proposal for a high altitude reconnaissance

aircraft. The request for proposal for the aircraft is

listed in appendix A. The performance requirements stated

in the RFP are listed below;

!.! REQUIREMENTS AND CONSTRAINTS

1. The cruise altitude will be 130,000 feet.

2. The required payload will be 3,000 pounds.

3. The design cruise Mach number will be M=O.7

4. The cruise (data sampling) time will be six hours.

5. There is a minimum of one crew member responsible

for piloting the aircraft.

6. A 6,000 mile range is required.
1.2 MISSION PROFILE

The mission profile is shown in figure !.I. The RFP

states that the total mission range is 6,000 feet. The RFP

also specifies that the sampling time, or time at cruise

altitude should be 6 hours. The range and time calculations

do not coincide with each other. Six hours at altitude at

0.7 Mach number correspond to a range of 3035 miles. If the

HI-BI was limited to six hours at cruise, then the climb and

descent legs would have to cover nearly 3,000 miles to make

up the rest of the 6,000 mile required range. Trade-off

studies using energy methods were performed to determine the

most efficient combination of climb, cruise, descent legs.

It was calculated that the optimum time at cruise should be

10.9 hours with a 5519.6 cruise range. The total mission

range meets the range requirement specified in the RFP but

exceeds the sampling time requirement by 4.9 hours.

2.0 INITIAL DESIGN

2.1 DESIGN CRITERIA

After rev_ewlng the aircraft requirements listed in

section I; certain assumptions about the aircraft were made.

One of these assumptions was that the aircraft would be

required to have a low wing loading due to the low dynamic

pressure at the 130,000 ft cruise altitude. Another

assumption made was that the aircraft would be propeller

3
driven due to the subsonic cruise requirement. Further, the

airplane would be required to have low drag due to the

difficulty in producing thrust at high altitudes.

2.2 CONFIGURATIONS

Possible configurations were considered for the

aircraft. These aircraft are listed in Figures 1.2 through

1.6. The flying wing in Figure 1.2 was considered for its

lack of horizontal tail and therefore its overall

aerodynamic efficiency. The aircraft would also provide a

large uninterrupted area to mount ozone sampling devices

(large leading edge). Further review of the flying wing

showed that the aircraft would require an expensive

stability augmentation system due to its inherent

instability. The aircraft would also require a long heavy

landing gear to accommodate the propellers and was therefore

rejected.

Two conventional monoplanes are shown in Figure, 1.3a

and 1.3b. A monoplane would be inherently more stable than

a flying wing and is a proven design configuration. The

preliminary weight and size estimations showed that the

airplane would have a wingspan of 600 feet due to the

aircraft's low wing loading. The conventional monoplanes

were ruled out due to the large required wingspan.

A canard was added to the monoplane as shown in Figure

1.4. The three surface configuration was considered for its

4
improved aerodynamic efficiency over the standard monoplane.

The configuration was ruled out however, because of the

interference effects of the canard on the main wing and the

destablizing effects of the canard configuration compared to

convectional area.

The joined wing configuration shown in Figure 1.5 was

seriously considered for the aircraft. The aircraft would

be more aerodynamically efficient than the conven:ional

monoplane but without the inherent instability found in the

flying wing configuration. The major drawback in

incorporating a joined wing would be that the production of

a joined wing aircraft would require the development of new

technologies as there are no large joined wing aircraft in

existence.

FIGURE 1.2

FLYING WING

AD_KNTAQEII

• HIGH AERODYNAMIC EFFICIENCY DUE TO THE

LACK OF A HOR|ZON'UU. TAIL

o LARGE FRON'I_L AREA R_JLABLE FOR

INSTALLING 8AMPLJNG DEVICE8

DI8ADtANT_E8

0 INHERENT INSTABIMTY

o LANDING GEAR/PItOP CLEARANCE

• POOR I_KEOFF RO170"ION

5

' ii !

I!:. !

?
The final configuration was the biplane shown in Figure

1.6. The biplane configurations maximizes planform area

while minimizing span. The biplane would thus have a

shorter wingspan then that of the monoplane while generating

the same lift. The configuration is inherently stable and

would have a more conventional structure than the joined

wing. The engines on the biplane could be mounted on the

top wing, thus solving the problem of propeller clearance

and landing gear size. The biplane configuration was thus

adopted for the high altitude aircraft.

High Altitude Biplane _

FIGURE 1.6

8
3.0 FINAL DESIGN

As stated in section 2; a biplane configuration was

chosen for the aircraft. The aircraft was thus named the

HI-BI, which stands for High altitude Biplane. The aircraft

has a twin boom fuselage to minimize the stress at the wing

roots. The aircraft has engines mounted on the top wing for

maximum propeller clearance. The engines would be aft

mounted with pusher propellers to assure uninterrupted flow

over the main wings. If tractor propellers were used in

front of the w_..g, _ the propeller wash would cause

aerodynamic interference. The aircraft would have a center

fuselage housing the pilot as well as the nose gear, and the

payload would be mounted in the lower wing for easy access.

For stability and control, the aircraft would have a single

horizontal tail _oining the two fuselage booms as well as

twin vertical tails. The HI-BI aircraft is shown in Figure

3.1.

3.1 PRELIMINARY WEIGHT ESTIMATION

In sizing the aircraft, an initial weight estimation

was made. Reference 6 was used to determine initial

aircraft weight. It uses weights of similar aircraft to

estimate a weight to begin the sizing of the aircraft.

Later, in the design stage, when the aircraft configuration

becomes more detailed, a refined weight estimation can be

made. The initial weight estimation for the HI-BI was as

follows:

9
TAKEOFFWEIGHT

FUEL WEIGHT

EMPTYWEIGHT

42000 Ibs

16079 Ibs

22921 ibs

3.2 WING GEOMETRY

The criteria for designing the wing is high lift and

low drag. Trade-off studies were performed to determine the

effect of wing geometry on lift. Figures 3.2 and 3.3 show

the effect of wing geometry on CLu" From these studies, the

wing was designed to have a high aspect ratio with little

sweep. The upper and lower wings have identical geometry.

The wing geometry is listed in table 3.3.

HI-BI RECONNAISSANCE AIRCRAFT

!t

, If

! i

m

FIGURE 3.1

10
6. Z5

Uariation of C-L-alpha with Uin_l _w_ep

{

6.Z8

6.15

6. le

6.95

6.0e

5,_5

5.cJ_

5.8'3

0

I I I I I I I I

2 4 6 8 10 IZ 14 16 2.8

_p at tic (deg)

Figure 3.2

Uariation of" Ming C-L-al_ha

_Ith nsl_C_ Rs_tO

6,0

5.e

u

4,@

3,e

Z.O

1.0 I ! I i I I I I I

@ Z 4 6 8 10 IZ 14 16 18 2@

_tZl_-l=t Ratto

Figure 3.3

Ii
From the constraint diagram of Figure 3.4, the wing

area per wing was determined to be 11086 ft 2 . This

corresponds to a wing loading of 1.73. The high aspect

ratio wing increases the wing's lift curve slopeand reduces

the drag due to lift. Because of the large wing area, the

aircraft does not need any additional lift during take off

and landing. Therefore, the wing has no high lift devices.

Ailerons are placed on the lower wing for roll control, and

spoilers are placed on the upper surface of the upper wing

to reduce the lift during landing.

3.3 HORIZONTAL TAIL

The horizontal tail was sized to trim the aircraft

during.cruise using the methodology of reference 6. The

elevator is used for pitch control. Table 3.3 lists the

horizontal tail geometry.

3.4 VERTICAL TAIL

The vertical tail was sized for longitudinal stability

and control using the methodology of reference 6. Table 3.3

lists the twin vertical tail's geometry.

-3.5 FUSELAGE

The fuselage for the HI-BI consists of a main center

fuselage and two fuselage booms. The center fuselage as

well as the booms are connected to the lower wing. The

12
dimension for the fuselage sections are listed in Table 3.1

and Table 3.2

Center Fuselage Dimensions

Average Diameter 3.8 ft

Maximum Diameter 5.9 ft

Body Length 45.3 ft

Body Side Area 170.05

Table 3.1

ftA2

Fuselage Boom Dimensions

Boom Diameter at the Tail 3.1 ft

Boom Diameter at the Wing 1.2 ft

Boom Length* 63.6 ft

* The boom length was measured to be the distance from

the wing trailing edge to the aft end of the boom.

Table 3.2

The diameter of the center fuselage at the cockpit area

is 4.9 which provides adequate space for a crew of one

person. The nose gear is also located in the center

fuselage, aft of the cockpit.

3.6 REFINED WEIGHT ESTIMATION

After sizing the wing, horizontal tail, vertical tails,

and fuselage for the HI-BI aircraft, the method in reference

13
6 was used to obtain a final weight for the aircraft.

final weight for the HI-BI was:

The

TAKE-OFF WEIGHT

FUEL WEIGHT

EMPTYWEIGHT

40622 ibs

14573 Ibs

22799 Ibs

Table 3.4 list the individual component weights for the

aircraft. The engine weight listed in the Table includes

the weight of the turbochargers.

HI-BI PLANFORM GEOMETRY

WINGS HORIZ. TAIL

AREA (ft "2 ) 11059

SPAN (ft) 471

ASPECT RATIO 20

ROOT CHORD (ft) 34

TIP CHORD (ft) 13

LEADING EDGE SWEEP 5

DIHEDRAL 0

c-bar 25

VERT. TAILS

TABLE 3.3

924 323

68 31

5 3

14 15

14 6

0 16

0

14 Ii

3.7 MOMENTS OF INERTIA

The moments of inertia for the HI-BI were determined

using the methodclog¥ of Reference 9. The moments of

inertia of the aircraft are as follows:

14
Ixx = 10800 slug-ft 2

Iyy = 10500 slug-ft 2

Izz = 12500 slug-ft 2

HI-BI COMPONENTWEIGHTS

COMPONENT LBS

WING - 2 5051 per

VERT. TAIL - 2 314.9 per

HORZ. TAIL 480.9

FUSELAGE& BOOMS 717.8

LANDING GEAR 1521.5

ENGINES - 3 1975 per

START SYSTEM 138.1

ENGINE CONTRCL SYSTEM 222.4

PROPELLER - 3 571.3 per

PROPELLER CONTROL SYSTEM 172.2

, FUEL 14573.3

FUEL SYSTEM 658.1

, ELECTRONICS i00

, INSTRUMENTATION 49.16

FURNISHING I00.I

AIR CONDITIONING 83.2 ,

CREW 250

PAYLOAD 3000

h

TABLE 3.4

15
3.8 CENTER OF GRAVITY LOCATION

Using Reference 6, the center of gravity location was

determined for the aircraft. Figure 3.5 shows a side view

of the HI-BI with the location of the center of gravity at

the take-off weight and at the empty weight. The center of

gravity travel between the two extreme conditions is .0747

times c-bar.

¥;w

CONSTRAINT DIAGRAM

HI - BI

FIGURE 3,4

HI-BI C.G. LOCATION AND TRAVEL

FIGURE 3.5

16

ORIG;NAL PAGE IS

OF POOR QUALITY
4.0 AERODYNAMICS

The geometric references used in

calculations for the HI-BI are as follows:

* Sre f = 11058.6 ft 2

MAC = 24.984 ft

* b = 470.88 ft

The reference area used for

calculations is the area of one wing.

the aerodynamic

the aerodynamic

4.1 AIRFOIL SELECTION

The criteria used for selection of the airfoil for the

HI-BI aircraft was as follows.

* low Reynolds number at altitude

* low drag at cruise

* low pitching moment at cruise

* high c I

Because of the low air density at cruise altitude, the

wing will be operating at a Reynolds number of 500000, and a

majority of the flow over the wing surface will be laminar.

For flow at low Reynolds numbers, a laminar separation

_'bubble" will develop on the upper leading surface of the

airfoil (Reference 3). If the bubble burst, the flow will

separate from the upper surface and lift will be lost. The

airfoil should be designed for low Reynolds number flow.

17
Figure 4.1 shows the variation of skin friction with

Reynolds number. Skin friction drag is inversely

proportional to Reynolds number. Because the HI-BI wing

will be cruising at a low Reynolds number, the zero lift

drag will increase. This is an undesirable quality because

it increases the powered required for the aircraft.

Low pitching moment is desired in order to decrease the

induced drag due by trimming the aircraft.

The airfoil selected to fulfill the requirements is

the Liebeck LNVI09A airfoil and is shown in Figure 4.2. The

LNVIO9A airfoil has the following characteristics (Reference

3):

* Clmax = 1.8

* Cla = 6.207

* Cmc/4 = -.05

* RN > 300000

per radian

The LNVI09A airfoil is designed for low Reynolds

number operation. The LNVlO9A airfoil also has relatively

constant drag coefficient over a large range of lift

coefficient. The airfoil drag coefficient is approximately

0.01 over a lift coefficient range of 0.4 to 1.35.

18
z....,

0

• ,01

TUR3R.JLEJqT S'KEIq FRICTIflSq COEFFICIEIqT

t.flS_"E5 REYJqGLDS _BEJt

8.4_01 ! ! ! !

1E_'4 1E'*05 1E-e6 1£-07 1E*4W

Juq

11"09

Figure 4. I

LNVIOgA AIRFOIL FOR HI-BI

f

I"

"---.......

1:3:<

Figure 4.2

19
4.2 LIFT

The lift for the HI-BI is generated by the biplane

wings. The lift of the biplane was determined using the

methodology in Reference 9, modified for a biplane. The

mission requirement dictated the aircraft was to cruise at a

constant altitude of 130,000 feet. At the beginning of the

cruise portion of the mission, the aircraft will be flying

at a total airplane lift coefficient of 1.53. At the end of

cruise the total airplane lift coefficient is I.I0.

Because of the close proximity of the two main wings,

there are interference effects between the two wings for the

biplane configuration. The lift generated by a wing in a

biplane configuration will be less than if the wing was in a

free stream by itself. Reference 7 shows the interference

effects between the two wings is negligible for a spacing of

the wings of .9 - i times the chord length or greater. The

wing spacing for the HI-BI aircraft is .9 times the root

chord.

The wings on biplane aircraft are usually staggered one

chord length with the upper wing forward of the lower wing.

This is done to prevent the lower wing from blocking the

flow of the upper wing during extremely high angles of

attack. Since the HI-BI aircraft will not be flying at

large angles of attack, a stagger of .08 times the mean

aerodynamic chord was used. The driving parameter for

2O
determination of the stagger was the placement of the c.g.

location to assure static stability.

4.3 DRAG

The difficulty of producing thrust at the required

cruise altitude of the HI-BI made it necessary for the drag

to be a minimum. The drag of the HI-BI was determined

using the methodology of Reference 6. Figure 4.3 shows the

drag polar for the HI-BI at the cruise altitude of 130000

feet. At the cruise altitude, the aircraft is operating at

a C d of 0.058. For the cruise C L of 1.53, the HI-BI is

operating at a lift to drag ratio of 26.4.

?

6 i

5 ............... "._............... 4. ...........

i

4

3

Z

DlUltnlGPIlL.MR FOR H I-BI

M -- 0,7 - IWl = 5eoeee

.............. 4 ............... _........... {................

CD

I

...... ? ............... _............... _ ............... ?

0.1 O.Z 0.3 0.4

Figure 4.3

21
5.0 STABILITY AND CONTROL

5.1 STATIC STABILITY

Static stability for the HI-BI aircraft was determined

using the methodology of Reference i0. The stability was

calculated for four flight conditions (phases) of the

mission profile, and the flight conditions are listed in

Table 5.1.

(LBS)

MISSION PHASE FOR STABILITY CALCULATIONS

PHASE ALTITUDE (FT) VELOCITY (FT/S) WEIGHT

1 0 50 40621.86

2 50K 140 39353.94

3 130K 742.56 38416.7

4 130K 742.56 27485.61

TABLE 5.1

Phase ! and 2 correspond to velocities for maximum

rates of climb during the ascent portion of the mission.

Phase 3 corresponds to the beginning of the cruise at

altitude, and phase 4 corresponds to the end of the cruise

at altitude. Table 5.2 lists the longitudinal stability

derivatives for the 4 flight phases. Table 5.3 list the

laterial-directiona! stability derivatives. The HI-BI

aircraft is statically stable at the four phases of the

flight mission (Reference 9).

22
LONGITUDINAL STABILITY DERIVATIVES

;'-.k3E

STABILITY " 2 3 &

DERIVATIVE

C-O-alpha O.5608 O.4565 O.9339 O.6682

i

C-O-u _ 0 0 ' 0

_,-L-_-h O.330:] O.3324 O.4058 O.4058

_-M-_-H -0.997 -1.0033 -1.224S -1. 1957

C-L-alpha 11.&296 II.5267 15.2943 15.2943

C-L-U O.0025 O.0213 1.4'/84 I.0578

C-L-q 8.3056 8.3734 l O.9726 8.77318

C-L-del ta-E O.1559 O.1579 O.1928 O.1928

C-L-alpha-dot 0.8!28 0.8251 1.3432 1.3111

C-m-alpha -I.&371 -I.4405 -I.4822 -0.381

C-m-q -7.9264 -7.9825 -I0.03471 -9.207778

C-m-delta-E -0.4736 -0.4755 -0.5818 -0.558

C-m-alpha-dot -2.4535 -2.4905 -4.0543 -3.8631

TABLE 5.2

5.2 DYNAMIC STABILITY

Dynamic stab111t7 was done for phase 3 and phase 4

using Reference 9 in conjunction with Reference 14. The

dynamic stability was done for these two conditions in order

to determine the effect the center of gravity travel had on

dynamlc stability. Table 5.4 shows the roots of the

characteristic equation for the longitudinal perturbation

equations of motion for the two phases. The aircraft is

dynamlcally stable about the longitudinal axis. Table 5.5

23
shows the roots of the characteristic equation for the

laterlal-directional perturbation equations of motion. The

HI-BI is also dynamically laterlal-directional!y stable.

LATERIAL-DIRECTIONAL STABILITY DERIVATIVES

phase

STABILITY 1 2 3 4

DERIVATIVE

C-y-beta -0.1361 -0.1361 -C.:361 -0.1361

C-y-p 0.0042 0.0038 0.3041 0.0036

C-y-r 0.0423 0.0425 0.3424 0.0415

C-y-delta-R 0.0281 0.0282 0.0339 0.0339

C-l-beta -0.0577 -0.0886 -0.1793 -0.1654

C-l-p -0.6135 -0.6194 -0.8195 -0.0195

C-l-r 0.3669 0.3669 0.3669 0.3669

C-l-deha-A 0.2135 0.2156 0.2989 0.2989

C-l-delta-R -0.0009 -0.0008 -G.O01 -0.0009

C-n-beta 0.021 0.0211 0.0211 0.0207

C-n-p -0.1725 -0.139 -0.2012 -0.144

C-n-r -0.053 -0.0442 -C.0669 -0.0475

C-n-delta-A -0.0132 -0.0107 -0.0229 -0.0154

C-n-delta-R -0.0044 -0.0045 -0.0054 -0.0052

TABLE 5.3

LONGITUDINAL ROOTS OF CHAR. EQU.

I PHASE 3 I

• S1,S2 = -.00295 +- J.08149 S3,S4 = -.92065 +- j2.88372

I

PHASE 4 I

I

I

TABLE 5.4

24
LATERIAL-DIRECTIONAL ROOTS OF CHAR. EQU.

PHASE 3

S1 = -.00344 $2 = -25.3165 $3,$4 = -2.542 _- j.68713

PHASE 4

SI,$2 = -1.752 _- ji.431 $3 = -.000275 $4 = -26.306

TABLE 5.5

5.3 FLYING QUALITIES

From Reference 5, MIL-F-8785B, the flying qualities for

the HI-BI aircraft can be evaluated. Table 5.6 list the

short period and phugoid frequency and damping ratio for

phase 3 and 4 for the aircraft. The flying qualities for

the short period and phugoid were determined to be level

two. This means the flying qualities are adequate to

accomplish the mission but with some increase in pilot work

load.

Table 5.7 list the spiral and roll time constants and

the Dutch roll frequency and damping ratio for aircraft for

the two flight phases. The flying qualities were found to

be level one for the two flights conditions. This means the

flying qualities are clearly adequate for the mission.

Loncitudinal Dynamic Response

zetaSP = .304

zetaSP = .304

PHASE 3

wnSP = 3.03 zetaph = .0362

PHASE 4

wnSP = 3.03 zetaph = .036

wnph = .082

wnph = .082

TABLE 5.6

25
Lateral Dynamic Response

Ts = .0395

Ts = .0380

PHASE 3

TR = 290.7 zetaD = .965

PHASE 4

TH = 3626.7 zetaD = .775

f

wnD = 2.63 l

!

wnD = 2.26i

TABLE 5.7

Figures 5.1 thru 5.6 show the aircraft's response to

unit step elevator deflections for the flight conditions of

phase 3 and phase 4. Comparing the plots for the two flight

conditions, the aircraft's response to elevator deflections

become less stable as the aircraft burns fuel (weight

decreases). The plane is still stable, but the time

responses increase and the overshoots become larger.

Figures 5.7 thru 5.10 show the aircraft's response to

unit step aileron deflections for the flight conditions of

phase 3 and phase 4. Comparing the plots for the two flight

conditions, the aircraft's response to aileron deflections

become less stable as the aircraft burns fuel. Again, the

plane is still stable, but the time responses increase and

the overshoots become larger.

26
VELOCITY CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

PHASE 3

0 i'

Q

, i

!IJ!\: ",! '"

lj Iv_

311 /_

FIGURE 5.1

ALPHA CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

PHASE 3

11211

I

FIGURE 5.2

PITCH ANGLE CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

HI

PHASE 3

I " "% .1%

: , , / _ _....'-'\._./_

f\,' ' ,,.],_.]'

' /

L

L

TI_ (_'_1

-5

FIGURE 5.3

27
VELOCITY CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

PHASE 4

15

If

I I

! , ,j \i

',/ v

Tilt (_.¢)

FIGURE 5.4

ALPHA CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

PHASE 4

-2

% //'X ,-% /_,

/ "1 , ", ,, ',. , ,.f_',, /'_"k#._",,_

I', " i' t k / "._ ./ _I_

'1'1_(see)

FIGURE 5.5

PITCH ANGLE CHANGE FOR UNIT STEP ELEVATOR DEFLECTION

PHASE 4

11 /'I ,,

Il

/] Ili ] I /I ti ,'",,:" '" ',.," -'

r , t..% .-

TIIrll ! ii I, '.. / ., J X,_f",

tl !/ 'ii ' / _"" '' '" "

I '

v

I I 1 a'

?sa

TII_ (s_)

FIGURE 5.6

28
BETA CHANGE FOR UNIT STEP AILERON DEFLECTION

PHASE 3

3

BETA

g

II L_ 50

TII_ (sec)

FIGURE 5.7

ROLL ANGLE CHANGE FOR UNIT STEP AILERON DEFLECTION

PHASE 3

jj

J

L0

hi

e |

|

lm

FIGURE 5.8

29
BETA CHANGE FOR UNIT STEP AILERON DEFLECTION

PHASE 3

iiI -_

5e

FIGURE 5.9

ROLL ANGLE CHANGE FOR UNIT STEP AILERON DEFLECTION

PHASE 3

2O

_[

0 Z5 5O

FIGURE 5.10

3O
6.0 PROPULSION SYST_

The propulsion system requirements, selection,

specifications, and performance will be described in the

following sections.

6.1 BYSTE4REQUIILm4ENTS

The mission profile for this aircraft sets very

stringent requirements for the propulsion system. At the

operational altitude of 130,000 ft., ambient pressure is

approximately 0.3% of standard sea level pressure. The

required 0.7 Mach cruise at altitude yields a condition of low

mass flow per unit area. In light of these conditions, the

powerplant for this aircraft must be able to operate without

large quantities of air, low specific air consumption. The

6,000 mile range requirement necessitates that the powerplants

have a low specific fuel consumption to reduce the amount and

weight of the fuel needed to complete the mission.

Since the aircraft is to operate at subsonic velocities

and very high altitudes, the aircraft's wings will be large

and heavy. This will require an engine that is capable of

producing large amounts of power at altitude. The final

requirements are to keep the engine and its systems as light

as possible and to develop this system with current

technology.

Figure 6.1 tabulates the requirements for the high

altitude propulsion system.

31
Figure 6.1

Requirements for

a Hiah Altitude powerplant

• Low Specific Air Consumption.

• Low Specific Fuel Consumption.

• Low System Weight.

• High Power Output.

• Utilization of Current Technology.

6.2 POWERPLANT SELECTION

Various engines were evaluated for their ability to

satisfy the requirements for a high altitude propulsion

system. The driving constraint in the engine selection process

was the air consumption of the engine at altitude. The air

consumption had to be low for the engine to produce power at

altitude. Figure 6.2 shows typical specific air consumption

(SAC} va|ues for the engines examined. The second constraint

was propulsion sTstem weight. The system weight is the weight

of the engine and weight of the fuel required for the

mission.This had to be kept as low as possible. Figures 6.3-

6.4 show typical specific fuel consumption (SFC) and specific

weight values for the engines examined. The engines evaluated

for this aircraft are described in the following sub-sections.

32
Figure 6.2

SPecific Air ConsumDUon

for Various Engine Types

SAC [lb_p-Ur)

Olml IlXat lB. Ioka'/ laltopeop lal_olot IKl_Zf_lnso

6

Figure 6.3

SDe¢111c Fuel Consumption

for Vccrious Engine Types

sFc t_,n_p-hr)

33 OF pOOR QUI'++.I.,.tTY
Figure 6.4

SPecific Weiaht

for Various Engine Ty'pes

Specific Wt. [lb/Hp]

6.2.1 TURBOJETS/TURBOFAHS

The use of turbojets or turbofans to complete the high

altitude mission was first evaluated. The Pratt & Whitney J75

turbojet, the engine used on the U-2 and the TR-1, was the

engine selected for examination. Figure 6.5 shows the

performance of the engine with altitude. The J75's thrust goes

from 17,000 lb. static sea level thrust to approximately 50

lb. of thrust at the cruise altitude of 130,000 ft.

The low density of the air at altitude and subsonic

cruise velocity combined with the engine's high specific air

consumption, make it impossible for any turbojet or turbofan

engine to produce any meaningful thrust.

34 ORIGINAL PAGE IS

OF POOR QUALITY
Figure 6.5

Thrust vs. Altitude

for a 17000 Ib Thrust Tarl_oJet

Thrust [lb/10"3)

20

16

1o

o s so ss 2o s6 40 so 6o 7o _o _o soo uo _o sso s40

Almude (ft/lO"3)

6.2.2 TURBOPROPS

After deciding that turbojets/turboprops were not

feasible, attention was turned towards turboprops. Turboprops

produce shaft power instead of accelerating air for thrust and

' have half the SAC of turbojets. However, turboprop engines

still require more air than is available at altitude.

Therefore, they follow the same power trend as the turbojet,

Figure 6.5, producing little power at altitude.

6.2.3 HYDRAZINE ENGINE

The hydrazine monopropellant reciprocating engine was

evaluated as a possible powerplant for the high altitude

aircraft. This type of engine uses hydrazine as a fuel and

does not require ambient air for combustion. The hydrazine

35
engine was developed by NASA for the Mini-Sniffer high

altitude aircraft, Reference 15. The engine for the MiniSniffer generated 15 Hp. An engine for this aircraft would be

a scaled up version of the Mini-Sniffer engine.

The hydrasine engine has an extremely high specific

fuel consumption, Figure 6.3, compared to other types of

engines. Hydrazine is also a toxic substance and must be

specially handled. Despite these drawbacks, the hydrazine

engine was considered for further study.

6.2.4 INTERNAL COMBUSTION ENGINES

Attention was given to exploring the feasibility of

using internal combustion (IC) engines for a high altitude

powerplant. IC engines have a relatively low SAC of 5-10

Ib/Hp-hr. The fuel consumption of these types of engines are

also attractively low, 0.3-0.5 Ib/Hp-hr. Although these

engines have a low SAC, they would be unable to produce enough

power at altitude without some type of supercharging. The

Lockheed HAARP Project, Reference 25, designed a turbocharging

system to operate with an IC engine at an altitude of 100,000

ft. It was felt that such a system could also be designed for

the required altitude of 130,000 ft.

A major drawback to IC engines is their high specific

weight, Figure 6.4. The high specific weight of these engines

added with the weight of the required turbocharging system

will result in a propulsion system would be extremely heavy.

36
Of the three IC engines examined, diesel, rotary, and

spark ignition, the spark ignition engine had the best mix of

SAC, SFC, and specific weight. The spark ignition IC engine

was selected for further study.

6.2.5 OTHER TTPES OF POWERPLANTS

Other engine technologies such as microwave propulsion,

laser propulsion, nuclear propulsion, and electrical

propulsion were examined. Practical versions of engines were

not feasible with present day technology and received no

further consideration.

6.2.6 SELECTION OF THE POWERPLANT

The two types of engines selected for further study

were the hydrazine monopropellant reciprocating engine and the

spark ignition reciprocating engine. Both engines were capable

of operating at the required altitude of 130,000 ft and

developing at least 500 Hp when scaled up. A system weight

study was conducted to determine which of the engines would

incur the least weight penalty completing a ten hour mission.

Figure 6.6 shows the results of this study with both engines

configured for operation at 130,000 ft.

The weight study showed that the spark ignition

propulsion system was four times lighter than the hydrazine

system. The main difference between the two engines is the

fuel required for the ten hour mission.

37
Figure 6.6

ComD,_'lson of Fatlmartmf Total

propulsion Welaht

S8

10-

16-

I0-

6-

0

To'k_ V_Jght ('Jb/lO'3)

Spm'k Iglflion H_Jrastne

man ila_ltsm Woi4jkt _ F_ml W_gltt

Norm: 10 hour Wtom

600 Hp onatno

hull_Od for 130,000 IL

Figure 6.7

Powerplant Conflaurartlon

• Reciprocmlng Spar_ lgnlUon.

• Horizontal Opposo(f Cylinders.

• Four Stage Turbocharged.

• Fuol Injected.

• Dual Ignition.

• Liquid Cooled.

• Geared Propeller Drive.

38
Thus, the spark ignition IC engine was selected as the best

choice for the high altitude propulsion system.

6.3 ENGINE CONFIGURATIOB

The configuration and details of the IC spark ignition

engine developed for this project will be set forth in the

following sub-sections. The engine configuration is shown in

Figure 6.7.

6.3.1

The high

turbocharging to

Turbocharging was

TURBOCHARGING SYSTD4

altitude engine uses four stages of

allow it to operate at altitude.

selected over supercharging so that the

engine power would not have to be used. Figure 6.8 shows a

schematic of the turbocharging system. Figure 6.9 tabulates

the specifications of the system. The turbochargers are each

composed of a radial compressor and a radial turbine. Each of

the four turbocharger stages are intercooled with a crossflow

air to air heat exchanger.

The full compression capacity of the system is only

required at the cruise altitude. The pressure in the system

is controlled by a waste gate installed between the engine and

the high pressure turbine, Figure 6.8. The waste gate is

designed to dump all exhaust up to a density altitude of 2800

ft. From 2800 ft. density altitude, the waste gate closes with

the decrease in density. Full closure of the waste gate occurs

39
at a density altitude of 97,000 ft. As an added safety

measure, an over pressure safety valve is incorporated between

the high pressure compressor and the engine.This valve will

release pressure if the pressure in the system becomes greater

than 2140 psf. This protects the engine from a potentially

disasterous over pressure from the turbochargers.

Figure 6.8

Schematic of the Four Stage

Turbocharging System

_P HP .P

stag. s,o_. slo_.I s_:_, r L7 __

IHL./ppp-- Law PreNm'v kderm_dla_e Pressure

WG - WoM. Got*

40
Figure 6.9

Specifications of the Four Stage

Turbocharuer System

TUrbocharger Type

Over All Pressure Ratio

lsl Stage Pressure Ratio

2nd Stage Pressure Ratio

3rd Stage Pressure Ratio

4th Stage Pressure Ratio

Maximum Mass Flow Rate

Maximum Pressure

Obtained at 130,000 ft.

Inlet Si_

System Weight

Radial

432:1

3:1

4:1

6:1

6:1

120.5 lib/rain)

1788 [psfa]

a.7(ft-2]

9oo (Lb)

6.3.2 ENGINE BLOCK ARD CYLINDERS

The high altitude engine is arranged in a horizontal

opposed configuration to reduce frontal area and allow an

aerodynamic cowling to be fitted around the engine. The block

is made up of two forged aluminum alloy pieces bolted together

vertically. The crank shaft is a forged steel, eight-throw,

one piece design and is supported by five journal bearings.

The engine has eight, 10:1 compression ratio, aluminum

alloy pistons displacing 1125 cubic inches. Each cylinder is

made up of aluminum structure with a forged steel bore sleeve

chrome plated to reduce wear and an aluminum alloy head. The

cylinders are bolted separately to the block allowing for

single cylinder replacement. There is one intake and one

41
exhaust valve per cylinder. The valve train is driven by a

single camshaft geared to the crank. The valves are connected

to the camshaft through a rocker arm-pushrod setup.

6.3.3 LUBRICATION SYSTDf

The lubrication system for the high altitude engine is

a pressure feed with a dry sump. The oil is pumped by a

positive displacement gear type pump and the full flow is

filtered. The oil receives cooling from a heat exchanger

mounted in the front of the engine cowling. The oil used by

the system is a 20/50 multi-grade.

6.3.4 COOLING SYSTI_

The cooling system used for the high altitude system

is a pressurized liquid system. The coolant used is a 60/40

mix of Ethylene Glycol and water pressurized to 14 psig. The

system uses a mechanical centrifugal pump capable of

delivering 125 gal/m/n of coolant to the engine. The system

has one radiator and a heat sink in the aircraft's fuel cell.

The mean temperature of the coolant is 210 F and maximum

system temperature is 265 F.

6.3.5 FUEL 8YST_

The fuel system for the high altitude powerplant

consists of a demand type mechanical pump with an electric

back-up pump. The fuel pump delivers the fuel to a electronic

42
metering pump. This pump will vary the fuel inputs to the

injectors to keep the correct air fuel ratio. There is one

injector per cylinder injecting the fuel into the cylinder

during the intake stroke. The fuel used by the system is 100

Low Lead aviation gasoline.

6.3.6 IGNITION SYST_

The ignition system for the engine consists of a dual

electronic ignition system. Each circuit is totally separate

and shielded with its own set of plug wires and plugs.

6.3.7 GEAR REDUCTION

A gear reduction box is employed to reduce the engine

RPM down to an acceptable speed for the propeller. The gear

reduction box is also used to mount and drive an auxiliary

alternator. The gear reduction box has provisions for mounting

other engine driven devices.

6.3.8 ELECTRICAL SYSTEM

The electrical system for the high altitude engine

consists of a single mechanical 24 volt alternator powering

dual 24 volt batteries and the engine sensors and aircraft

systems. An extra alternator is driven by the engine to supply

power for the payload package.

43
6.3.9 ENGINE CONTROL STSTEMS

The engine control system for the high altitude engine

is split in two parts, pilot controls and computer controls.

The pilot controls the engine through dual throttles and

pitch levers. The computer controls the engine's fuel mixture,

ignition timing, and turbocharger waste gate to achieve the

optimum performance.

The information on condition of the engine is displayed

to the pilot through RPM gages, oil pressure gages, cylinder

head temperatures, coolant temperature, and manifold pressure

gages. Any engine faults are recorded by the computer for

later retrieval.

6.4 PERFORMANCE SPECIFICATIONS

This section describes the performance of spark

ignition engine developed for this aircraft.

The powerplant was modeled on an engine program

modified from Reference 19. The program takes design

parameters for the engine, preforms a cycle analysis, and

outputs the performance of the designed engine. The program

simulated the an eight cylinder engine and turbochargers

operating at 130,000 ft. Figure 6.10 shows the specifications

and performance for the engine designed for this aircraft.

Figures 6.11 and 6.12 show cycle information on the engine's

pressure vs. volume and heat transfer vs. gas temperature

respectively.

44
Figure 6.10

Performance Specifications

960 Hp Engine

Engine Type

Number of Cylinders

Cylinder Arrangement

Born and Stroke

Displacement

Compression Ratio

Width and Height, Engine

Width and Helght, InstaLled

Length and Frontal _ Engine

Length and Frontal Area, lnst

Engine Weight

Tolal Weight, Installed

Weight/Horsepower

Fuel Grade

SFC, Cruise and Max Power

SAC, Cruise and Max Power

Cruise Power

Max Power

IC Spazk Ignition

8

Horizontal Opposod

5,25 in and 6,5 in

1125 cu in

10:1

38 in and 29.25 in

41 in and 59.8 in

33,6 in and 7.7 _ It

69.6 In and 16.4 sq ft

11177 lb

2077 tb

1.89 lb/Hp

100 LL

0.357 and 0.383 lb/Hp-hr

5.684 and 5.45 lb/Hp--hr

962 Hp/3900 RPM O ]30k fL

1194.9 Hp/4250RPM • SJ_

II00 Hp/4250 RPM • 130k It.

45
Figure 6.11

Pre___.ure w. Volume Dlaomm

960 Hp Engine

_ure (i_st)

i 800"

1000 -

SO0 "

0

0 40 60 80 108 _0 140

Volume [tn'31

Clulse Concltttont

Dtspla_ment • 11_ cu.t_

FigUre 6.12

Heat Tran_R!or vs. Temperature

960 Hp Engine

I I I

i II 1 • I

l_mperatum/(lO"3) (_

K_ltne Dtsplacemat - U_ ctt.in.

46
6.5 PROPELLER DESIGN

After analyzing the mission profile, the cruise phase

was determined to be the flight phase governing the

propeller size. This is due to the fact that the air

density is very low at the cruise altitude of 130,000 feet.

The two main criteria at cruise for the propellers were:

* Since the air density is low at altitude, the

propeller wlll have to have a large diameter

* The tip velocities can not exceed the local

sonic speed because of compressibility effects.

Crl'LIDal Tip _l_l_t"i_

Figure 6.13

47
Figure 6.13 is a graph of constant tip velocity for M =

1 and M = 0.8 as a function of angular velocity, n, and

propeller diameter, D. Any combination of n and D above the

M = ! curve will give tip velocities greater than the local

sonic speed. At this altitude, the sonic speed is 1060 fps.

In fact, the tip speeds should not exceed a Mach number of

0.8 because local velocities on the blades can go sonic.

Therefore, The propeller should operate in the region below

the M = 0.8 curve. As can be seen in the figure, this

region is very narrow for propeller diameters of over 20

feet.

It was determined that at cruise, the drag was equal to

1708 pounds, _hich for a constant velocity, is equal to the

thrust required. This value of thrust corresponds to 2304

horsepower for a cruise velocity of 742 fps (M = 0.7).

Therefore, for the three propellers, each must produce 768

hp.

48
7.0 PERFORMANCE

7.1 TAKEOFF

The take-off performance for the HI-BI was evaluated

using methods outlined in References 4 and 6. The wing

loading on takeoff was found to be 3.86 Ibs/ft'2 and with

the air density and CLmax known , the stalling velocity was

determined to be 30.9 fps. This is a very low value for

stall velocity due to the low wing loading. Therefore, it

was decided to stall the top wing with spoilers in order to

increase the wing loading and thus increase Vstall. In so

doing, the aircraft will not lose lift due to wind qust.

Vstall was increased to 43.7 fps. The take-off velocity Vto

is then 1.2 times Vsta!l and has a value of 52.4 fps. The

acceleration is assumed to be constant and taken at .7 times

Vto. A force balance on the aircraft was done to get the

acceleration. The distance traveled by the aircraft from a

zero velocity to Vto was then calculated to be 390 feet.

This is the ground roll distance SG. The distance in which

the plane then rotates into take-off position is known as

the rotation distance SR and was calculated to be 157.2

feet. Next, the transition distance STR was determined.

This is the ground distance that the plane actually travels

as it climbs through a constant velocity arc. This distance

was equal to 70 ft. and at this point the plane has not

cleared an imaginary 50 foot high wall. The distance to

49
clear this barrier was determined to be 372.3 feet.

Therefore, the total take-off distance is equal to 989.5

feet.

When talking about take-off performance, one needs to

look at balanced field length. This is the distance needed

for the aircraft either to take-off and clear the barrier or

brake and come to a stop in case of an engine failure.

Bala_cd Fi"ld L_th

Fig. 7.1

The distance is said to be balanced if at a given velocity,

the total distance traveled up through takeoff is equal to

the distance, at the same velocity, traveled with one engine

5O
inoperative and the aircraft is brought to rest. From

Figure 7.1, the balanced field length is the distance

corresponding to the intersection of the two curves. This

is read off the plot to be about 1750 feet. The

corresponding critical speed, Vcrit, is 43 fps. Therefore,

if one engine fails at a velocity higher than 45 fps, the

take-off should be continued. If the failure occurs at a

lower velocity than Vcrit, The take-off should be aborted.

7.2 Landing

For the landing analysis,

22748 ibs and a CLmax of 3.6.

the weight was equal to

Thus, Vsta!l was calculated

to be 31 fps. The total landing distance can be thought of

the sum of three parts, the Air Distance SA, the Free Roll

Distance SFR, and the Braking Distance SB. Using methods

outline in Reference 6, SA was found to be 325 feet. This

is the horizontal distance that the plane travels after

having passed over the 50 foot wall up through touch down.

The Free Roll is the distance traveled as the plane's nose

pitches down and the front landing gear makes contact. This

distance was equal to 106.5 feet. The distance to a

complete stop, SB, was found using a braking coefficient of

.5. This distance was determined to be 32.3 feet. Thus,

the total distance for landing was found to be 463.8 feet.

Table 7.1 summarizes the values obtained for the various

distances.

51
TAKE-OFF W/S=3.86 T/W=. 199

T/W=0

Ground roll SG 390 ft.

ft.

Rotation SR 157.2 ft.

ft.

Transition STR 70 ft.

ft.

Climb SCL 372.3 ft.

ft.

LANDING W/S=2.05

Air SA 325

Free Roll SFR 106.5

Braking SB 32.3

TOTAL 989.5 ft. TOTAL 463.8

Table 7.I

7.3 CLIMB

The climb performance was determined using the methods

of Reference 12. The objective of the climb analysis was to

minimize the time the HI-BI was required to climb to the

specified cruise altitude. Minimizing the time to climb

minimized the fuel required to climb to cruise altitude.

Figure 7.2 shows plots of the HI-BI's rate of climb as a

function of airspeed and altitude. To minimize the time

required to climb, the HI-BI's velocity should correspond to

the maximum rate of climb at the corresponding altitude.

Figure 7.3 shows the maximum rates of climb as a function of

velocity taken from Figure 7.2. Using the maximum rates of

climb from Figure 7.3, the time to climb to 130000 feet

cruise altitude was determined to be 66.2 minutes. Using

52
the time required to climb, the fuel required to climb to

the cruise altitude was determined to be 1330.75 !bs.

7.4 CRUISE

The RFP stated the aircraft was required to either

cruise at a 130000 foot altitude for 6 hours or to fly a

total mission distance of 6000 miles. The constraining

factor to determine the cruise distance was the 6000 miles

total mission distance. The HI-BI was determined to cruise

at altitude for 10.9 hours. A Mach 0.7, this corresponds to

a cruise distance of 5519.6 miles.

w--

CE

Rate of climb as a function

of airspeed and altitude

4°'°°!_ ' /

34.88]_F_'i ...............T..............i..............i...............i..............i..............1

_-___' .............. ........ _, ............... :,............... ,:............... ,:.............. t

_°_°t'_........ !t_i_ ........ i\i.... i..... %_! ............. -.............. 1

__t_....... _11_.............. Y_ ............... i...... %i.............. i...... -...... 1

4.13 _ i i i............._..............1

I L I |

-- 1 _ll_ _1_ I ...... , i J

.... 0 100 200 300 400 500 600 700 800

velocity (ft/s)

4- h=OK -e- h=2_Sk -e- h=50k

--- h=T4k --- h=lOOk 4- h=130k

FIGURE 7.2

53
dle.ee

_1.00 q

1@ ,_'

_ximm rate or c l lmb

function of al itude

r •

i i z .:

i i " i

.................... _.................... _....................................... ._............... 4.......................................

......... ,o,o, ...... i ........ ,o, ......

1

i I i

altitude (ft)

14@I@@

FIGURE 7.3

7.5 F11ght Envelope

HI-BI FLI_T ElqUELOPE

1__[ 140000

1Z_ l

@ .

• 1_ Z_ 3@@ 4_ 5Q0 _ ?_0 8@@ 9@@ 1@00

VELIX:ITY (ft_)

Figure 7.4

54
The flight envelope in Figure 7.4 was developed using

methods outline in Reference 4. A flight envelope is the

area on an altitude versus velocity plot where the aircraft

can fly. At low speeds, the aircraft is constrained by the

stall velocity and at higher speeds the constraint is

maximum power. The top curve on the Figure represents stall

velocity with increasing altitude. The stall velocities are

related to altitude by the following equation:

(Vstall) 2 = (2W/S)/(rhOoCLmaxO)

The density ratio o varies with altitude thus at any given

altitude there is a unique Vstal I. From Figure ?.4, it can

be seen that Vstal I is about 30 fps at sea level and gets to

be about 500 fps at 130000 feet altitude. The stall

velocity at 130000 feet is well below the operating velocity

of 742 fps.

The maximum power curve also increases with altitude

and follows the stall curve closely. This gives a flight

envelope that is somewhat narrow. Thus, much care must be

taken when piloting HI-BI in order to stay within the flight

envelope.

7.6 POWER REQUIRED

Using the drag determined earlier and Reference 6, the

power required for the flight mission was determined.

Figure 7.5 shows the HI-BI's required power as a function of

55
velocity and altitude. The Figure shows as the plane

increases in velocity and altitude, the power required

increases. The cause of the increase in power is the

decrease in Reynolds number with increase in altitude. As

stated earlier, the drag of the aircraft increases with

decrease in Reynolds number. Figure 7.6 shows a plot of

Reynolds number as a function of altitude. From Figure 7.5,

the power required to cruise at altitude was determined to

be 2'68_ .7 hp.

L_

O

Powered

6000

4500,

4000,

8500,

8000,

2600 _

2000,

I 600,

1000,

600,

°olo

_ 0

-i- 74k

Required For Various Altitudes

20 80 40 50 60 70 80 {aO

velocity (f/s)

--- 25k -e- 60k

4- lOOk 4- 180k

10

FIGURE 7.5

56
Uaria'tioY, oF i_y'mnmlds #

wl'_h Rlti_ude - U : 74Z fps

FIGURE 7.6

8.0 STRUCTURES

The aircraft was assumed to be composed of three types

@

of structural elements. The three elements are stiffened

shells, stiffened plates, and beams. The center fuselage as

well as the twin booms are stiffened shells. The top and

bottom of main wing stiffened plates. The wing itself is

supported by beams extending from the fuselage and ending at

the wing tips.

57
8.1 STRUCTURALANALYSIS

A static load analysis was performed to determine the

structural integrity of the main wing for the preliminary

design. Two flight conditions were simulated in this

analysis; normal flight, n = 1 g, and flight experiencing

the maximum gust load factor, n = 6.8 g.

8.2

The

section.

of spars,

FINITE ELEMENT ANALYSIS

wing planform was modeled for the half-wing

The structural arrangement of the wing consisted

ribs, and skin stiffeners. Spar members are

continuous along the wing span and form a torque-box (wing

box) carrying the main load component. The ribs are

segmented elements that help stabilize the skin elements.

The wing structure was modeled on computer using the

MSC/NASTRAN software package. A finite e]_.. ,-nt analysis of

the model was performed to optimize the distribution of

stress on the internal structure of the wing. The wing skin

material used in the analysis was graphite epoxy with the

following properties_ El=3Oxl06psi, E2=O.75xlO6psi,

v12=0.25, and G12=O.375xlO6psi. The stacking sequences for

the angle-ply symmetrical laminates were 0, 45, and 90

_egrees. The thickness of the wing skin was 0.5 inches.

The wing was discritized into 5 spars and 25 ribs, 9.81 feet

apart. The upper and lower surface of each rib has !0 nodes

(spar-rib junctions). Thus the entire wing has 260 nodes

58
and over 780 degrees of freedom since each node has 3

degrees of freedom. This network of spars and ribs is

covered with a machined graphite epoxy skin on both the top

and bottom surfaces. Graphite expoxy laminate was chosen

for the wing model because of its relative high static

strength, long fatigue life, and resistance to corrosion.

Moreover, graphite epoxy can be fabricated for a lower cost

when compared to similar metallic structures. The finite

element model of the wing is given in Figure 8.1. The wing

was modeled in a local rectangular coordinate system. The

local coordinate system for the wing has the X-axls positive

outboard, the Y-axis positive toward the leading edge, and

the Z-axis positive up based on the right-hand rule.

8.3 CONSTRAINTS

The constraints on the model were created to counter

any unbalanced rotation on a wing. Since spanwise

elliptical lift distribution was the main loading on the HIBI wing model, only minimal amounts of constraints were

applied. The constraint points were put along ribs were the

wing model was attached to the fuselage. All three

translational and rotational degrees of freedom were

constrained at these wing root nodal points. The nodal

points along the rest of the wing were free to translate and

rotate in the X< Y, and Z direction.

59
0 0

h

4

E

!

m

u

o

w

¢Nb

o

.°

o _

o

g EVe

_ Oow

4_

o_

X

Figure 8.1

60 ORIGINAL PAGE !S

OF POOR QUALITY
8,4 LOADING OF T_E MODEL

A V-n maneuver and V-n gust diagram was constructed for

the aircraft for two flight conditions, Figures 8.2 thru

8.3. The flight conditions considered were; steady level

flight at sea level and steady level flight at the cruise

altitude. The positive load factors at cruise are 3.5 and -

!.2. The critical load factors were shown to occur at sea

level where they 6.8 and -3.7.

Once the critical load factors were determined, the

wing spanwise elliptical lift distribution of the

appropriate load can be obtained by the following:

L' = L' o (i - (y/(b/2/))2)I/2

where, y...Spanwise location along the x axis

measured from the centerllne of the wing.

L'o...Lift per unit span at the center of the

wing.

b...Wing span

The results of the applied loads analysis are tabulated

in Table 8.1. The wing spanwise elliptical lift

distribution accounts only for aerodynamic loading. The

total load on the wing structure is composed of the

aerodynamic load, the structural weight load, the fuel

weight load, and the engine weight load. The aerodynamic

loads are much greater than the weight loads and act in a

direction opposite the weight loads. The net effect of

including the weight loads in the analysis is that the

overall structural load is reduced.

61
V-m I_kll_UV[E Ol&GIrJU'l FAI)_

3.0

2.4

1.0

!

8

- 1.0

-2.0

ClUISIll Sill[E| CONOITIOm

VD - _lil DIVINi LI_[O - Ill.S4 ill

VC " {_'SI_i CIUISI_I _l_[Dz 43g,il ill

VA - KSISN _A_EU_[Ilmi Sli_EO • 439,61 ill

V$ - STALL SPE[0 AT *11 ,* 388.61 ill

¥0

SPIED, V

ktOo *'AS

Figure 8.2

V-n MANEUVER DIAGRAM FAR 2_

SEA LEVEL CONDITION

VS • 24.PJ Irtll

4 VA • 28.62 ktl

VC • 38.4S ktl

VO • 48.07 ktl

. , b-il.'il I...............

. i_v,

| 'oL----_'_._ 3'o

J

-2

Y_gure 8.3

62

VC

• I

:4O

._ SPEEDoV

kts. [AS
¥-sGIJ_lrOlaKIAR FAR25

3.0

2.0

1.0

i°

!

-I.0

VD

NJ SPEIE0, V

tte o I[AS

CRUISIIm8 $11_ED CONDrT|Oll

VO - D_SIEli DIVINI $PEn • alg.54 ktl

V_ - OE_SI_ CRUISIIm SPfI[D : 439.61 ttO

VII - IIESIBR SPEED F01t HAEIMUR

tlrrERSITT, 414._ kt8

Figure 8.4

TABLE I: WING LOAD DISTRIBOTIOH (n : It)

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

i

t

i

L

¢

A

LIFT/ PLOAD1 PS, OAD1 PLOAD1 PLOAD1 PLOAD1 WING tOIL

b/2 (lb/f¢) (lb/ft) (lb/ft) (lb/ft) (lb/ft) I_o/x_; IAu, ,

-;:0_- _8;.0, ,.31 45., ,.88 _,_,; 21.21 -21.41-31.38

,.81 169.,8 ,.28 ,.90 3,.54 2,.,3 ,1.1, -1,.--28.28

19.62 169.30

29.43 188.66

39.24 167.52

4g.06 166.1?

58.86 184.50

88.67 182.51

78.48 180.18

88.29 157.50

98.10 154.44

107.91 181.00

117.72 147.13

127.53 142.81

137.34 137.99

147.18 132.82

156.96 126.63

168.77 110.92

176.58 112.37

186.39 103.80

196.20 93.91

208.01 82.25

215.82 8T.gO

225.83 48.53

235.44 0,00

35.18 46.78 37.55 29.65 21.14 -17.99 -26.77

35.03 45.58 37.38 29.52 21.06 -16.39 -24.39

34.81 46.30 37.15 29.34 20.92 -14.87 -22.12

34.53 . 44.93 36.65 29.10 20.75 -13.42 -19.96

34.19 44.48 36.48 28,81 20.54 -12.04 -17.92

33.77 43.98 36.04 28.48 20.29 -10.74 -15.98

33.29 43.32 36.52 28.05 20.00 -8.51 -14.16

32.?3 42.59 34.93 27.58 19.67 -8.38 -12.44

32,10 41.76 34.25 27.06 19.28 -7.28 -10.84

31.38 40.83 33.49 26.44 18.85 -8.28 -9.35

30.58 36.?9 32.63 25.7_ 18.37 -5.35 -7.96

29.68 38,82 31.87 25.01 17.83 -4.50 -6.69

28.68 37.32 30.60 24.17 17.23 -3.72 -6.53

27.56 35.86 29,41 23.23 16.56 -3,01 -4.46

26.32 34.24 28.08 22.18 16.81 -2.38 -3.54

24,92 32.43 26.60 21.00 14.97 -1.82 -2.71

23.35 30.39 24.92 19.68 14.08 -1.34 -1.99

21.57 28.07 23.02 18.18 12.66 -0.65 -1.38

19.52 25.40 20.83 16.45 ll.T3 -0.50 -0.88

17.09 22.24 18.24 14.40 10.27 -0.33 -0.50

14,11 18.35 15.06 11.89 8.48 -0.15 -0.22

10.09 13.12 10.76 8.50 8.06 -0.04 -0.08

0.00 0.00 0.00 0.00 0.00 0.00 0.00

63

ORIGINAL PAGE IS

OF POOR QUALITY
8.5 DEFLECTION

Figure 8.5 and 8.6 show the wing deflection in the x, y

and z directions. The deflections in the x and y direction

under a 1.0 g load were 3.42 x I0_-2 feet and -2.9 x I0"-2

feet, respectively. The displacement in the positive z

direction was determined to be 2.84 feet. At the critical

load factor of 6.8 g, the maximum deflections in the x and y

directions were -2.16 x I0--I feet and -1.75 x i0"-I feet,

respectively. The maximum deflection in the z direction was

15.7 feet.

8.6 TSAI-HILL FAILURE CRITERIA

The Tsai-Hill theory was used to determine if the

graphite epoxy would fail in the NASTRAN finite element

program. The Tsa!-Hil! equation yielded a failure index

between zero and 0.894. The failure indices were below one

which indicates that the graphite epoxy will not fail.

8.7 STRESS

The minor and major principal stresses under the normal

flight condition of ! g, were found to be -1.59 ksi and 1.26

ksi, respective. The limit load factor of 6.8 g applied to

the aircraft yielded minor and major stresses of -!0.83 ksi

and 7.43 ksi, respectively.

64
DEFLECTION DUE TO LIFT

DEFLECTION (it)

18

iI ....

0.00 39.24 78.48 117.72

SPANWISE LOCATION (it)

Normal Flight lg "-4---N - 6.89 I

Figure 8.5

I

DEFLECTION DUE TO DRAG

-0.1

-0.12

-0.14

DEFLECTION (It)

L ' ' : ' ' ' : ' ' ' I,' ' ' ' ' ' ' ' ' ' ,

0.00 39.24 78.48 U7.72 156.96 19630 238,44

SPANWISE LOCATION (ft)

I -- Normal Flight, lg -4-- Limit Load, 6.8g

Maximum Wing Deflection:

n - lg, Y--.008ft

n - 6.82, Y • -.122 ft

Figure 8.6

65
8.8 BENDING MOMENT

The finite element analysis showed that for the normal

flight condition of 1 g, the maximum bending moment occurred

at the 4th spar. The bending moments for this point are

listed below:

Normal Flight (n = Ig)

Mz = 1.42 x I0"4 ft-lb

My = -1.04 x 10"4 ft-lb

The bending moments for the limit load factor of 6.8g

were determined for the wing. The results are listed below:

Limit Load Factor (n = 6.8 g)

Mz = 1.2 x 10"5 ft-lb

My = 5.08 x 10-4 ft-lb

9.0 LANDING GEAR

9.1 LANDING GEAR ARRANGEMENT

The HI-BI's gear is a tricycle configuration and is

shown in Figure 9.1. The nose gear is 25.3 feet in front of

the center of gravity (c.g.) with a forward swept angle of

41 degrees. Also, the nose gear structure is tilted 7.5

degrees backward. On the other hand, the main gear is 5.9

feet behind the c.g. with an aft swept angle of 15.0

degrees. Looking at a bottom view of the HI-BI (Figure

9.2), the main gear is 34.1 feet from the aircrafts

66
LANDING GEAR PLACEMENT

W,R.T. CENTER OF GRAVITY

113,4 ft.

_A

ft

16.0 ft.

T

6.0 ft.

7.0 ft.

Figure 9.1

DOWN AND LOrXF.D

l=-_l

37.4 iff.

Figure 9.2

6'7
centerline, and is 37.4 feet from the nose. The landing

gear employs a floating link system with the main retracting

forward into the wings and the nose retracting rearward into

the fuselage.

9.2 RUNWAY LOADS

Static load analysis was performed on the gears to

determine the normal forces acting on the runway (per

Reference !I). Alowing for a 25% airplane growth, the

following loads resulted:

I) A main gear static load equal to 10864.7 ibs.

2) A nose gear static load equal to 5401.8 Ibs.

3) A nose gear dynamic load equal to 12569.8 ibs.

9.3 LANDING GEAR TIRES

The tire selections were based on compatability with

large well maintained runways. Table 9.1 shows the tires

chosen for the HI-BI.

9.4 SHOCK ABSORBERS

Each gear has a single oleo-pneumatic strut with an

axle fastened directly to the strut piston. The main gear

shock absorbers are 30.4 inches long by 6 inches in diameter

and each are capable of absorbing 560,532.5 ft-lbs of

kinetic energy. Conversely, the nose gear shock absorber

measures 39.0 inches by 6.2 inches and withstands 568,490.0

ft-lbs of knietic energy.

68
SELECTEDTIRES

MAIN:

• TYPE VII

• TIRE O.D. -

• MAX. WIDTH =

• UNLOADED INFLATION

PRESSURE =

• MAX. LOADING = J.L.9_0.fl_LL_

• WEIGHT PER TIRE =

NOSE:

* TYPE VH

• TIRE O.D. =

• MAX. WIDTH =

• MAX. LOADING =

• UNLOADED INFLATION

PRESSURE =

• WEIGHT PER TIRE =

Table 9.1

!

GEAR ASSY

TOP

TO_OUE

LINK

SIDE BACK

Figure 9.3

69
9.5 GEAR STRUCTURE

The gear structure was scaled from an aircraft with

approximately the same take off weight. The geometry is

shown in Figure 9.3. It's made of steel alloy and weighs

about 1600 lbs. The main gears are fastened to spars

located in the booms, and the nose gear is fastened to the

spars in the fuselage.

9.6 LANDING GEAR HYDRAULIC SYSTEM

The HI-BI utilizes a state-of-the-art light weight 8000

psi primary system with titanium lines. The working fluid is

Ch!orotrifluoroethy!ene (CTFE), a nonflammable fluid. It's

fastened to the HI-BI's fuselage. An emergency pneudraullc

back up system is also located in the fuselage. The

redundant system generates 3000 psi manually.

9.7 BRAKE SYSTEM

Each main gear has an independent braking system. The

system uses CTFE at 8000 psi through its titanium fluid

lines. Anti-skid carbon brakes are installed in each main

wheel. It features a five rotor brake with an oversized

insulating ring. Each brake is self-bleeding and contains a

temperature sensor. The pilot controlls each brake

seperately. To illustrate, the pilot depresses the right

rudder toe pedal to actuate the right-hand wheel, and

depresses the left toe rudder to actuate the left-hand

wheel.

7O
!0.0 RELIABILITY AND MAINTAINABILITY

The reliability of the aircraft was determined using

statistical data compiled at the Northrop corporation. The

data that was looked at was the MTBF (mean time between

failure) and MTBM (mean time between maintenance). Based on

these data, the operating hours, sorties, MTBM and MTBF of

the !I major systems, the corresponding systems on the HI-BI

aircraft were given a complexity or simplicity factor based

on the particular design and component. It was determined

that 82% of all failures are inherent and 18% are induced.

Further 45% of all maintenance is inherent, 10% induced and

45% are no defects maintenance.

Eleven main systems were looked at in the reliability

analysis; namely the airframe, fuselage, landing gear,

flight control, electric power supply, lighting system, fuel

system, instrumentation, radio and engines. Then average

mission duration for the HI-BI was calculated to be 13

hours. Each component was given a reliability percentage

based on the compiled data for each component. The initial

statistical analysis considered strictly a non-redundant

series configuration. This configuration assumes that if

any individual component of the aircraft fails, then the

whole aircraft would be unoperational. Using this concept,

the reliability of the aircraft after 14 hours of flight

came out to be approximately 68%, which is not a good

71
overall reliability. Thus, the non-redundant approach

proved to be overly conservative and unrealististic.

A more realistic approach was then taken in the

statistical analysis. This approach considers only the

major systems; namely the airframe, fuselage, engines, fuel

system, and flight controls. This approach yielded a

reliability percentage of 78 _ for the 13 hour mission,

shown in Figure I0.I.

HOURS VS. RELIABILITY

•I. Itm

0 10 20

HOURS

FIGURE 10.1

72

OF PCOR C_JALIIY
The maintainability is also a major part of the

project. In Figure 10.2, the eleven systems have been

illustrated along with their maintainability constants.

This constant explains the number of times the system needs

maintenance before failure, therefore the lower the

maintainability constant the better off the system is. As

it one can see in figure 10.2, the main areas of concern

would be the engines, the fuselage and the airframe.

In order to calculate the mission availability and equipment

availability, a maintainability and availability analysis

program was used to calculate these variables for each hour

of operation of each system. The program and output can be

reviewed in appendix C.

SYSTEM VS. WUUNYAMAIIUl_

_me

m_UUM_

M. 0OWSL

L$ IJ IJ

_d

IU_TL_J_JTT

FIGURE 10.2

73

ORIGINAL PAGE IS

OF POOR QUALITY
II.0 COCKPIT VISION AND HUMAN FACTOR

The cockpit was designed to provide the pilot with the

standard field of vision 15 degrees below the line of sight

as shown in Figure II.i. The cockpit also had to

accommodate the standard pilot. A study was performed to

determine the average size of a male pilot. Body component

weights are for a male pilot with a weight of 179.3 ibs. are

as following:

BODY COMPONENT WEIGHT IN LBS

Head and neck 15

Upper torso 49

Lower torso 28

Upper legs 39.9

Lower legs and feet

Upper arms

Lower arms and hands

29.8

9.9

7.7

1"/9.3

The figures for the dimensions and weights would have

to be 85% of male pilot for a female pilot. It was decided

that we can not deviate from this particular weight and

dimension due to the nature of the aircraft design.

The figure 11.2 illustrates how the exact geometry and

configurations of the human body and how it is implemented

in the cockpit design of figure II.i.

74
PILOT LINE OF SIGHT

....... _I_Wal-_ _ _. _ _ - [

' ¢k

r [ _._,/ .,_[d.; _ ;" ,.

L: -

FIGURE 11.1

HUMAN GEOMETRY

, _4M.__

._5 .... _ ___ 'I._.

_ ,. _..,.:.o .

!,' _, ' . _

"._."-/_ _

' I,', iT{

i 0 , _ p:mT

. t.--_L

FIGURE 11.2

?5 ORIGINAL PAGE IS

OF POOR QUALITY
!2.0 COST ANALYSIS

In order to come up with an estimate of the cost,

historical data had to be looked into to utilize the

regression met_._ds of forecasting more efficiently. The

cost of manufacturing, tooling and engineering were looked

at. Also relationships were examined that would explain

values of time put into different aspects of the aircraft.

The variables that proved to have the greatest explanatory

values were AMPR weight, maximum speed and the quantity.

Exponential regression models have been used to fit the best

statistical curves. Incidentally, these equations provide

coefficient of correlation from 83% to 97% which is good for

estimating cost. In fact, most contractors use the same

ideology to estimate the expenses. Some of the breakdowns

are as following:

I. Engineering: By using these exponential

equations, engineering hours could be estimated,

which could then be used to determine engineering

cost.

2. Development support: This cost encompasses

the cost of manufacturing labor and material

required to produce mock-ups, test parts, static

test items and other items of hardware.

3. Flight test operations: This category

includes performance reliability, stability,

control characteristics and air worthiness etc.

this also includes engineering planning and data

reduction, manufacturing support, instrumentation,

76 OR_;G:,",c,'.. PAGV. IS

OF POOR QL_ALITY
spares, fuel, oil and pilots' pay, facilities

rental.

4. Tooling: Tooling hours are defined as hours

charged for tool design, tool planning, tool

fabrication, production test equipment, check out

of tools, maintenance of tooling, normal changes

and production planning.

5. Manufacturing labor hours: This includes

labor hours necessary to machine, process,

fabricate and assemble the major structure and to

install purchased parts, administration and off

site manufacturing assemblies. Based on this, the

cost will be forecasted.

6. Manufacturing materials: This includes the

raw material, hardware and purchase of parts

required for the fabrication and assembly of the

major structure of an aircraft.

7. Quality control: quality control is closely

related to direct manufacturing labor, but is

estimate separately, since records of actual hours

for quality control are not normally included in

general manufacturing hours data. On the average,

it amounts to approximately 13% of the total labor

hour.

The following are the project constraints:

I. The AMPR weight:

2. The s_eed of aircraft:

3. The :rctotype quantity:

4. The nreduction quantity:

5. The due date of production:

6. The value of learning curve:

25000 ibs.

400 kn

I

4

1998

.8

77

ORIGINAL PAGE IS

OF pOOR QUALITY
The learning curve of 0.8 is standard in the aerospace

industry and based on the 0.8 learning curve it was decided

to produce 4 aircraft. Also due to inflation and the nature

of the equations used, a program was written to convert the

dollar values to the value of 1998. So all the dollars

values are based on the final completion of the project. In

order to come up with a flexible way of computing the

values, a program was written and is shown in appendix B.

LEARNING CURVE

OUANTTI_

FIGURE 12.1

12

78

ORIG_NA_ PAGE IS

OF POOR QUALITY
The following is the breakdown of the cost. The final

cost per aircraft was determined using the standard 0.8

learning curve.

Production development cost:

Production quality control cost:

Production manufacturing labor cost:

Production test flight operation cost:

Production engineering cost:

Production tooling cost:

Production total cost:

Production cost per aircraft:

Learning curve value per aircraft:

$34,015 730

$12,960 270

$54,632 060

$15,162 880

$42,920 540

$45,223 450

$349,832,800

$78,481,090

$25,889,100

!3.0 MANUFACTURABILITY

The decision was made to manufacture four production

aircraft and one prototype. The production of these

aircraft during the span of 8 years requires certain number

of employees in tooling, engineering and manufacturing.

Assuming 50% allocated to R&D and 50% allocated to

production and assuming 2000 working hours allocated to each

employee. The following figures were obtained.

Total number of employees in engineering: 179

Total number of employees in tooling: 270

Total number of employees in manufacturing: 350

799

79
14.0 ELECTRICAL SYSTEM

The HI-BI has four independent electrical systems. The

electrical power - under normal operating conditions - is

aquired by engine driven generators. When reversed, these

direct current (DC) generators double as starter motors.

The primary power is inverted to alternating current (AC)

via DC buses.

Each system is carefully shielded from the effects of

lightning. And the crucial buses and wiring bundles are

widely seperated. Additionally, all the batteries are

shieled from the primary structure and flammable materials.

15.0 LIFE SUPPORT

The pilot will be required to be in a space suit

becuase of the low pressure at altitude. Enough oxygen will

be carried to complete the 14 hour mission. Even though the

pilot will be in a heated space suit, the cabin will be

required to be heated to heat the flight instrumentaion.

16.0 CONCLUSION

Although the intent of the design group was to design

an aircraft that would be capable of achieving the

performance outlined in the RFP, several problems have

presented themselves which indicate that the mission cannot

be performed with current technology. The problem areas are

listed below:

8O
Aerodynamics

The horizontal tail may be too small to generate enough

downforce to rotate the aircraft at takeoff. The

airfoil was selected for its performance at the cruise

Reynolds number rather than its performance at the

cruise Mach number. The airfoil might not be able to

provide the lift that the aircraft requires.

Propulsion

Supercharging system has yet to be installed on an

engine of this size. The actual operating efficlencies

may be much lower than estimated because the

calculations were single stage units rather than series

arrangements. A propeller has yet to be developed that

is large enough to produce the thrust required. The

problem of cooling the engine has to be solved as the

traditional method of convection would not be feasible

due to the low dynamic pressures found at cruise.

Structures

The limit load factor did not include a factor of

safety. The airqraft would probably be required to

withstand higher load factors because it would be

manned. Further the structural analysis performed on

the aircraft considered only the main wings and the

supports. The overall structure will probably need to

be redesigned to withstand the high load factors

experienced by the aircraft.

81
Human Factors

The long flight duration of 14 hours is probably too

demanding on a pilot. The ER-2 has a similar working

environment and pilot workload to the HI-BI. Pilots

have much difficulty in withstanding flights longer

than 8 hours in the ER-2.

17.0 RECOMMENDATIONS

The 130,000 ft. cruise requirement was the most

demanding performance requirement. The cruise altitude

should be lowered to an altitude closer to i00,000 feet. At

this altitude the dynamic pressure is four times greater for

the design Mach number of 0.7.

The 6 hour cruise time also seemed unrealistic. This

required the aircraft to carry a large amount of fuel given

the difficulty in producing thrust at altitude. Scientist

indicate that the long cruise time is required to sample a

large section of the ozone layer. An alternative to the

long cruise time might be to fly the aircraft at an even

lower altitude and climbing to the desired altitude for very

brief periods thus sampling a large section of the ozone

layer.

82
REFERENCES

I ,

°

,

°

o

.

°

°

°

IO.

II.

!2.

13.

Abbott, I. H. and Doenhoff, A. E. Von, "THEORY OF WING

SECTIONS", Dover Publications, Inc., 1959

Houghton, E. __. and Carruthers, N. B. , "Aerodynamics

for Engineering Students," Edward Arnold Ltd., 1982

Liebeck, R.H. and Camacho, P.P.; "Airfoil Design at Low

Reynolds number with Constrained pitching moment";

Douglas Aircraft Company; Long Beach, Ca

McCormick, B. W., "Aerodynamics, Aeronautics, and

Flight Mechanics," John Wiley and Sons, 1979

MIL-F-8785B (ASG) Military Specification - Flying

Qualities of Piloted Airplanes, August 1969

Nicolai, L.M.; "Fundamentals of AIRCRAFT DESIGN";

Aerospace Engineering University of Dayton;

Dayton, Ohio, 1984

Olson, C.E. and Selberg, B.P.; "Experimental

Determination of Improved Aerodynamic

Characteristics Utilizing Biplane Wing Configuration";

AIAA paper, University of Missouri- Rolla, Rolla,

Missouri, 1974

O'Neil, J.D.; "Maintainability and Availability

Program", California State Polytechnic University

Pomona, Pomona, Ca.

Roskam, Jan; "AIRPLANE FLIGHT DYNAMICS AND AUTOMATIC

FLIGHT CONTROLS"; University of Kansas; Lawrence,

Kansas, 1982

Roskam, Jan: "Methods For Estimating Stability And

Control Derivatives Of Conventional Subsonic

Airplanes"; University of Kansas; Lawrence, Kansas,

1982

Roskam, Jan; "AIRPLANE DESIGN PART 4", University of

Kansas; Lawrence, Kansas, 1982

Shevell, R. S., "Fundamentals of Flight," PrenticeHall,Inc., 1983

Vail, J.T., "Design for Maintainability", SAE Technical

Paper

!4. AFDA, VERSION 3.2, Aladaire International Inc., 1989

83
15.

16.

17.

18.

19.

20.

21.

22.

23.

24.

25.

Akkerman, J.W., Hydrazine Monoprope!!ant Reciprocating

Engine Development. NASA Lyndon B. Johnson Space

Center. Houston, Texas, January 1976.

Lockheed-California Co., "Stress Concenteration Factors

Reference

Logan, Daryl L., "A First Course in the Finite Element

Method", PWS Dr_ub_ishers, _ Boston, MA ., 1986

Schaeffer, Harry G., "MSC/NASTRAN Primer: Static and

Normal Modes Analysis", Schaeffer Analysis INC., Mont

Vernon, NH

Benson, R.S. and Whitehouse, N.D., Internal Combustion

Engines. Pergamon Press Inc. New York, 1979.

Kays, W.M. and London, A.L., Compact Heat Exchangers,

3rd Edition. McGraw-Hill Book Company. New York, 1984.

Zinner, K., Supercharging of Internal Combustion

Engines. Springer-Verlag. New York, 1978.

Janota, M.S. and Watson, N., Turbocharging the Internal

Combustion Engine. John Wiley & Sons Inc. New York,

1982.

Campbell, Ashley S., Thermodynamic Analysis of

Combustion Engines. John Wiley & Sons Inc. New York,

1979.

Benway, Ralph B., Design an Development of a Light

Weight, High Pressure Ratio Aircraft Turbocharger. SAE

871041, 1987.

High Altitude Atmospheric Research Platform,

Information Package. The Lockheed Corporation. Burbank,

California. Febuary, 1990.

84
APPENDIX A

REQUEST FOR PROPOSAL
DESIGN OBJECTIVES AND REQUIREMENTS

REQUESTFOR PROPOSAL: A HIGH ALTITUDE

RECONNAISSANCEAIRCRAFT

I. OPPORTUNITYDESCRIPTION

The recent discovery of an ozone hole above the North Pole

has prompted the scientific community to accelerate their

efforts in investigating man's impact on his environment.

The existence of the ozone hole has brought about concern

that the predictions of environmentalists may come true.

Unless the ozone depletion in the Earth's atmosphere is

controlled, radiation levels at the surface may increase to

harmful levels. At the equator, the ozone layer ranges from

80,000 to I00,000+ feet which is beyond the capabilities of

the ER-2 aircraft, NASA's current high altitude

reconnaissance aircraft. Therefore, to effectively

investigate the ozone layer, NASA needs to develop a high

altitude airlift which will reach altitudes of I00,000+

feet. To hasten the development of the technology and

methodology required to develop an aircraft that can reach

these altitudes, the NASA program has been working closely

with industry and universities. Perhaps, with the data

retrieved by this aircraft, scientists and politicians will

be able to formulate an emissions control plan which will

diminish the rate of degeneration of the ozone layer.

II. PROJECT OBJECTIVE

The objective of this project is to develop three possible

designs for an aircraft that can cruise at I00,00+ feet and

sample the chemistry of the ozone layer at this altitude. 1

An innovative approach which pushes the limits of existing

technology is inherent in the nature of this project.

However, the techniques Used should be feasible by the year

1990. All operational constraints must be met.

III. REQUIREMENTS AND CONSTRAINTS

I. The cruise altitude will be 130,00 feet.

2. The required payload will be 3,000 pounds.

3. The design cruise Mach number will be M=0.7

4. The cruise idata sampling) time will be six hours

5. There is a minimum of one crew member responsible

for piloting the aircraft.

6. A 6,000 mile range is required.
APPENDIX B

COST BREAKDOWN
"1" ,'•" ",./r. - ,- ........................... '"

THI:-; PRNGRAH I.,.!ILL FIGURE NUT THE TIME ANC: I-:FLST A_" Ei'4GIr,!EERING, Trlr;Lrr,.;_3

_..4L., r..tANLIFAF:TLIF.:iF.4:J OF _ r"-"-=.,_ ...... _,_,i _ PE _i R CF.:_FT _,N[:, TH-,_ F-'FODLii-:TI Oi_ C,F Erie

' PARTICULAR H!R CRAFT,

•"******************************t****A*[*** OUT LINE*******************************

C L S

GF_'SUE_ INPI_ITDATA

GO'SUE: PROTFz

GOSUE: PRO[:'

G0'.:-;1_1E',C,EFLATOR

G0 :--_; U B C 0 ST

i_U.:,L,B P RI"-,IT

E N D

iNF'UTE:,ATA :

INPiJT "ENTER THE AI".IF'F.: t,..iEiGHT OF TEH AIR CRAFT";

INPUT "ENTER ]-HE SPEE[:, OF AIR CRAFT ,::1-:::I.._:;"; :_::

INF'UT "ENTER THE PROTOT,PE _-:.!UAI.]TIT",""; C4P

:NPLIT "ENTEE' THE F'ROE:,IJFT!F_!.._ F!UANTIT""I C!

INPUT "ENTER NC. rEAF.::--; AL::}C:ATED FOR THE F:_:F-i..:E:Z:T: ; F:

INPUT "THE PEF.:CEHT LEARi'-i!I.!G" _ N

R ET U RN

................ .-e._ .,...,- ..,-..,-..,-, .***** _ _******* _** _**** _ F'F.:I-IT L-_T T E,.*.._**.:,: .-e:_** _._,.*******************

P R 0 T 0 :

_HbINEERING

LET EF' = ,::S.,634 * _ ,.=,7/.:, * ('::; .:-:56::, * ,::QF' .96::,

__PRINT "THE TOTAL F'ROTOTYPE EI.iF_-.iNEERING HCU_':--: _RE": EP; "HOURS"

:'ROTOT","PE C,EVELOPMENT '.:;LIFPORT

LET E:,F' = ,. ,0&474 * A " .:-:.' ._.... :,. ,. S 2..,'_.. ' * F!F .4:-::5::,

F'R 0 T ',ZlT"," P E T Fi 0 L I N G

LET T;' = =-._,,, ::::::5 * A ,_-,-",,z.. . ..':--. ' . 6:-:.-:: _ ,"L-..!F' .482::,

LF'_:I/4T "THE T!ZIT,.:4L PROTOT,'FE TOCL!!4G HFli_IF..:::_,qR:E: : TP; 'HF!I_IF.::E:''

" !::'£OTOT"CPE H_i"..tlJFACTUR I t".]G

LET LP = ' .:-:OI'T' "* ,q " i . ! 1',-:'., _ , ':; .41 .:, * .: L:.!F' I .:-:._..-_,::,

LF'RIf.- = "THE TqTAL F-'RF_TOT'_F p [iAI.4UFACTJR.f.I.iG HZHj_:; _F..:E" ; L=; "HOI_!FL:-Z:

:_'OTOTYFE HATERIAL

LET PtF = (! .:5 * A " .5F-_',5::, * , .: i ,.':.'!:-;::, ÷ L:!P " ..-L,22 .

"F'ROTZITYF'E L-.!LIAL:T"r ' ;-:OHTROL

LET CQCP = ,13 * LP

' PROTOTYPE FL I GHT TE:=;T

LET F F' = < . 0 !:i 1244 _ A 1 . 16., ÷ ,::S " I . 371 _ ,: C!F.... i , 281

R ET U RI"4

F ROD :

ENG INEER I NG ORIGINAL PAGE IS LET E = ' 0.... "_ + ..........

...... _,.-..... i_.:.:_, OF POOR QUALITY

LF'RINT "THE TOTAL ENGIr,_EEF.:ING H_-'_UR:S ARE"; E; "HOURS"

" C,EVELOFt.IEI'..IT SUPPORT

LET [:, = (.(:108325 * A

FLIGHT 3EST OPERATION

LET F = ,001244 *

TOOL I r.4G

'ET T =

,_PRINT "THE TOTAL TOOLING HOURS ARE"; T; "HOLIRS"

"MA[.4LIFACTIJRING HOURS

LET L = '::28.':?'.---:4 ÷ _ .74, + , ':: .54:-:::, * ,::g': ..=,24.,

LP'Ri r-.,, ._T ....THE T,qT_L I.iAi'..ii_iF_:-:Tl_!E: i i_3 HOLI_::-, ARE ; L ; ' HAl _ ..... IE' :; ,,

["I_I"._I_IFAiZ:T!_IE:Ibig COST

"r .. ? C' "--' ,
E:ETuRN

" ********************************* DE**,A**************************************

DEFLATOR :

AER']SF'ACE CPI FOR I770 I:-; 35.. ?

ET "FR = l'_"?O + R

.-:HL:-:JLATIi'-.IG THE [:,EFLATOP

LET Li = YI_:- 196_:,

LET CF = lO.?l:E:d,:': + '::4,Sf;':E:93::-; * LI::'

i_ CF- THEN

Z,:.: = ,:.CP .... 35,7) ,.:1 . (2El + R>:., - 1

F'IF = I00 * IR

Ebi [:, I=

F.:ET U F t.-A

.C-!-3T :

i,: IF:: THEfi

-,: = L.' * ',1 + IR::' '::21:! * R

[:'F't = C'P + ! + IR) ':.20 + R'

P!P': = HF' + : + IR) " <2Fr + P,

Zi = :: i+ ': i -_ !IS:::, " '::28 -_ tz

,:-: i = v p ÷ t + I R:;, ' ,::2n + R .:,

ENA : _.EER I I.._G,•,r.IAr-._L!FAC:TLiF I i.!G _--"-; '- ' ' i..... __Trii3 H,_II_,I_.Lr _:HTES

uET T1 = Y_: - 1'..;'54

;F T i = YR - 1954 THEN

_E = 5.45::: + .=,.=,,q* T I

Y +'. = r-; . 1 :_::4 ÷ , ._-,-' z" 7',-', * T i

"-,T = 5.748S + .340,5 + TI

q[:, I'--

uET --' "-_-I = EF' +" "(_--

;__E-T :--'I = TF' + 'Y'T

::- : h_ = E + "(E

LE 7 TT = T a,. ,"T

_ET r-1 = H + Yt"t

__ r- i ;_ __ t ÷ "fJ'-."

_E- _F-- = '(Pl * LF

_E"- izL-!i-_= .1:':; * LF'I

LE7 CC!C = .i:-'.+ Hi

LET _ = [',F'! + MFI + EF'I + Fm'- - LPI + CQCF' + TF'I

LET E: = FI -_ E: + TT + f"tI _" TT + CQC + DI + LI

! ='= TOT = A + B + :::<P

LET .::::: = TOT .. (QP + Q)

LET >..::_:'= .1 * TOT

'LEARNING CURUE

T:::::: = :::;:: +.... gt N

R E T U Rhi

r--'F.,+p..JT:

F'RINT "THE AHPR t,..IEIGHT E= AtRF:RAFT" I A O_l_i.%,.f_LIp_ ,lS

PRIt-._T "TH = ::PEEr. OF AIRCRAFT ,.KI.'}"; :- OF POOR QUAL!T'/

F-R!r.;T "THE NE:, OF PROTOT"-_"PES REQUIF.:ED"I QP

-_II._T "THE NO, OF PRODLi'.Z:TIOI.._ REC!:JtRED" ; O

RIP..]-" "THE NO, OF "(EAR.':; ALUOL-:ATED FOR PROJECT"; R

F't::I_4- "THE I/.4:=sATIOf.._ RATE B_SE[:, ON 1770 FOR"; YR; " I'.-:" ; "1.." ; PIR

FR]I-.i- "THE E:,UE [:,ATE FOR THE F_.OJECT IS"; '_R

F'lq:',l4- "THB PROTF;T'yFE [:,EUELOF'I.!E[.!T COST IS"; [:,Fi ; "[:,O_.__ARS"

F+RIi._T "THE F'ROTOTYF'E rIATERiA__ s-:O:_::- i':;", MF'I I "[:,OLL_I_.::-;"

_-F-:IN- "THE _-'R_-ITO-","FE L::.!UA_IT_ i-:L-If.ITR:!-:L i::iJ::;l ,''-_:'' i-:l_-..!_[:F" "[-:'Oi..___RS"

F=RINT "THE PROTOT,"PE TE:-_-;-_ FLIGHT O;'E_:_TIOP..I CO'.--:. iS": F;- : "F:,NLLARE. '=
APPENDIX C

MAINTAINABILITY & AVAILABILITY ANALYSIS
150 PRINT

1 _0 PRINT

i70 PRINT

130 PRINT

190 PRINT

200 PRINT

2 I 0 PRI_;T

220 PRINT

230 F'EINT

2&0 PRINT

250 PEI_rT

260 PRIHT

I00 COLOR 7, 9, t: CLS

110 REM: J.D. O'Neil 5/85

120 PRINT " NAINTAINABILITY & AVAILABILITY ANALYSIS"

I30 PRINT

i,_0 PF:INT "Th g pro,_r'a:Ti 3F_2lyzes ,4 number of different maintenance and a.vai

par'ameter's foi' a system. The user' is asked to provide a number ¢

:i:,ecifi[,:atior, s anti the time Tnter vat fo_' the analy'_--is."

The specifi.:atfons required ar'e:"

!) The mean time of a maintenance _ct_on (phi)"

2) The mair_tenance time constraint (t)"

3) The system failure r'ate (lambda or r)"

,_ TI-,_ m[:.s!on ._tar't t_me {usually -0) u

S) Th'- _ ler_,qth of time for' the __r,al'tsis"

6) The time _r_cr'ement for data analysis"

NOTE -- .,_1! times should be in hours."

2_0..... PF:It4T "Th- {.:r_-.,-i_'am.= prov_de_, the f_]la_in._, o,,tput:".,

"30 PRINT

.gO PF,!NT u,, _',',.I_ITAIN,_.3][I TTY -" d=f[r;es ,_ *he _ !D!",'-_bilitv _!., .... t"'t t'a _.--+,_=In ,.

-- -- i l ,-1 e m "/ /_,7 L i

300 ........... It;Tr'F. ";_:+; =._ t.". '-,i"= ........ : :' * ional =-_:*'-. +-i'ver:es_ w_tl--Tn :: .-vet." pei io ,-!_ c +',

- _ L " ,.. L..... "

i:D FF:I_!T ": ,..= .r2 r_te.:-,arLLe t_me 'l.'Oi-KS_ 3lr.._' w_J=_, the nta_r_terL3nce ._,z_-_c,i:

........R20 PF,[HT "fci'nl=_J i ......... -,-!'d3i_ce_ k,_it _-I: i:_ escr';'- -d_.,_ pi'ocedutes. .

• 30 PRINT

2%0 INpIIT '_ ENTER to con,_ , ride , ,_$ : CLS

350 PRINT "(_) HAINTAINABIL1TY 1.4._REMEt4T - defines the propor't_on of fa_lu

360 PRINT "the rnissior, which w,]l be Iestored to oper'at_onal effectiveness

370 PRINT "mainten_r, ce time constr'aint (t}: .as a result cf maintenance .act

280 PRINT

• -L . 3.90 PRINT ", -"> _'.-,g,_IPHEftT._ _.VAI' A£,ILITY. is the pr'ob.=_il{fv., that_ a stated p=

L.00 F'RINT "o- equipment w_ll be_ avail-'-le=,_ for use at a g_v-n- t{me. due to +I::.

_t_lO PRINT "l:.(n-_{_, effect ,of the survivors an,J t:, units restored to ser'v[ce t:

z_20 PRIFT "n,_,r_t_.-Ti_.sL[l-it.:. fn a time equal to o_ less that the mainter, ance -+

::,30 FK:!NT HO'.)il&tI'.3;i-,t, "

L,t_O PF:INT

';_0 PFZINT ',",','! t,l!S:IC, ii ',.'AIL.A.BTLITY [_ the r-,Iob_.b_l"ty that a stated F--rc-

:_CO PRINT 'm; zzic:P-- _z _ve +_ d_ _-_ w _'' _ -

. ,_ ; _ _ r.. . .me ,, _=,_,or, , _ I nc,. h=ve ar,y f21]_Jr'e 114 ._:.:

.:_0 PRINT '_mi_=z;i,:,,_ wl-:ic!_ ,.:anr, ot, thi ou.:jh m_.ir, tain_bflity, be !e_to:'e:l to se

L.GO F'RINT " _: _ t_n,-z lez-. *:hart or equal to tl:e rnL_ir, ter_ance time constla_r,t,'

_90 PRINT

_SO0 PRINT " 5) UP-TIME-RATIO (UTR) - the ;atio oi: the time the system is '-.

510 PRINT "oper'atinci to the sum of the down and up t_me. It _s composed of :

520 PRINT " ......

..... =,.= element (mu/(r+mu)) and e transient element in the early oF530 PRINT "per'_od, UTF:+DTR=I . "

5_0 PRINT

550 INPUT "/ENTER>. to confil4u_"_ , Z$

560 CLS

ST0 R$ = "#######.######,,

590 E = 2.7 I {3281828#

310 INPUT "Mean time of mairrten._=.nce action (phi)? " PH

620 MU = _, .," PH

6_0 INPUT "',.#h.at {s the mair,+enar, - ce _ime - con_=traint 't),,, ? " TC

66,0 B = MU _" TC

350 INPUT "',#hat is the system failure rate (failures/hour')? ", R

660 INPUT "What is ttl__ star'tin,_ time of the ar, alys{ -_; "_ T:

670 INPUT "What is the ending time for' the arealisis? " TE

680 INPIjT ..... "At wha+ intervals do yo,, want to see +'n-_- ar, a 1,/s [ s _: " ; D_

690 CLS

700 PRINT "Mean m._ir, te,q_nce action t_rn_ (hr's)="; PH

710 F'RIN- "Mainten3nce _.,:tions per hour -"" 1 ' PH

PRINT '_M._ir, tenance time :onsti'aint (t} ='; TC

720 F'RINT "S/stem f,_lure r ate =": R Of POOR O,UALiT'_
" _< /F,I ''_ _'_.,__1i F,4f i _ _ ' -£' ....

_50 F'F-:Z,NT

'60 f,c:I,qT 'Miss:,or,, Maii_tei_a,nce Equ(pm-:i_t Mi_io(_

770 F'RI;_T P' HOOI'S UTP Increment Availabi]it,{ Ava_l_b_lTty

730 FOR T = T3 TO TE 3TF_P DT

790 RMIJ = R _ MU: IF T = 0 GOTO 810

800 TRM = T • (RMU ^ 2): MRT = R / TRM

810 IF T = O THEN UTR = 1 ELSE UTR = (MU / RMU) * MRT MRT m E * (-(RMU '_

'$20 HD -- 1 E "" (-R * T) (E " -B) _ (1 - E ^ (-R _< T))

830 LE = 1 - (E (-B)) _ <1 - E " (-R * T))

8b,O LM = E ^ ((-R * T) * E - (`-B))

850 F.T = E ^ (-R * T)

860 PRZi'IT USING "If####'; T;

870 PRINT USING R$; UTR;

880 PRINT USING R$; MC,;

890 PRINT USING R_; LE;

900 PRINT US]]NG R$; LM;

910 PRINT US1ZNG R$; RT

920 NEXT T

1000 END

ORiGiNAL PAGE IS

OF POOR QUALITY
Mean maintenance action time (hrs)= 182.2

_la_ntenance actions p_r' hour = 5._.88474E-03

Naint=nance time constraint (t) = &0

1.900007

0,995O15

0.990099

0 9£_252

0.980471

O.975757

0.£71107

0.960522

S,OO3C0C

0 001969

0 O03919

0 005848

0 007759

0 009651

0 011523

0 013377

,.000006

0 991979

0 98z_079

0 978178

6 968395

0 900690

0 953062

0 945510

009000

0 991971

O 984007

0 97610?

0 968270

0 96O496

0 952785

0 945135

, 000000

0 990010

0 980120

0 970329

0 960636

0 951039

0 9_1539

J 932133

_',,-,ar_ ,naintenance ruction time (hi's)= 132

,'i.e. ir, tenance actions per' hour' .: 7,975758E-03

lla_ntei-,ance time col',_:tr'-:int

1.000000

0.995249

0.960170

0.985879

0.£81L_56

0.977007

0.972630

0.968324

0 000000

0 00124C

_2 C82689

0 003982

0 0'1!5286

0 006576

0 00785a

0 099119

t) = 20

! OOO000

0 991306

] 9:33691

0 975652

0 907691

0 959805

0 95199a

0 94_258

0

0

0

0

0

0

! 000000

{ 99180!

982663

975603

967604

£59670

951802

9_3998

1.000000

0.99046E

0.981022

0.97i669

0.962_05

0.953229

0.944141

0.935139

,"_ean m.mintenance action time hr's)= 160.3

_4-_ntenance a,ct_ons per hOClF : 6,238303E-03

ha;:,tenance tlme .:onst_._fnt (t) = 8

1.000000

O 999796

0 98'9659

0 98a506

0 939629

0 974727

0 969899

0 965143

0.09{'000 1.000000

0.000508 0,990063

0.0Oi012 0.980231

0.001510 0,970501

0.002002 0.960872

0.002490 0.95t3&4

0.002972 0.941916

0.003450 '0.932586

I O00000

0 990061

C 980221

0 970_78

0 980832

0 951283

0 9&1828

0 932467

i 000000

0 939555

0 9?9219

0 968991

0 958870

0 948854

0 938944

0 929136

Mean maintenance action time l-,r's;= 135

b_aintenance actions per" hour = 7.407407E-03

_la_P, tenance time constr'a_nt (tl = 8

!.000000

0 994309

0 988836

0 983397

0 978053

0 97279£

0 967636

6 962551

0.000000

0.000650

0.001293

0.001928

0.002556

0.003177

0.003791

0.00a398

1.000000

0 98835_

0 9?8829

0 968422

0 958133

0 9&7960

0 937902

0 £27958

1 000000

0 989351

0 978815

0 968391

O 958079

0 9&7876

0 937?82

0 9 __,795

1 000000

0 98870_

0 977536

0 96649a

0 955577

0 944783

0 93Z_111

0 923560

ORIGINAL PAGE [_;

OF POOR QUALITY
Mean maintenance act

Mair, ter, ance actions

Maintenance time con

1 .000000

0.996235

0.992520

C.988855

0.985238

0.981669

0.978196

0.97/4671

ion

per" hour' =

str'aint (t)

0 000000

0 000368

0 00079Z_

0 001097

0 001457

0 001314

0 002168

_q O_q_

time (hr's)= 160

.00625

= 8

! , 000000

0.992817

0.985088

0 . 9700 ; 3

0.971592

0 .964623

O,957707

0 .95094 3

I GO0000

0 992816

0 985693

0 978601

0 971571

0 964591

0 957661

0.950780

1.000000

0.992449

0,98z_954

0.977517

0.970135

0.962809

0.955539

0.948323

Meai, m.=,intena_,ce action tim_ {hrs): 2390

[4,9i nter, ance __c±_onz per houi : :_.18_# 101E-O&

M::;_,t_:r,ance t _me constraint :t : : 8

i , 000000 0 , 900000 ! , 000000

0.q_99_,_00 t-:,O0,_, _'OL:i L ,__9.°801

:2:.999800 0.00000'1 0,99:3601

0,999700 0.000002 0.999q02

0 . 999601 0 . 000003 0 , 92:9203

0. 99950 I C. 000003 O. 99900_

0,999_# 01 O. 00000_, 0. 998805

0 . 999302 0 . 00000E, 0. 998606

:.090000

C.999803

0.999601

0.999402

0.99920]

0.99900&

0.398805

C.998606

1.000000

0.999800

0.999800

0.999400

._9200

0.999000

0,998801

0,998601

Meaii ffl,9 i AtecLance _ct

Haii-ten_nce actions

_.'t9l iqteJ:ance tirn< :on

1,000000

0.997787

0,995598

0.993_02

0.99129O

0.939072

0.986927

0.984790

p ,_i h o u r

st :',_, _ nt

0

0

0

o

0

0

0

0

time :':,r s) &02.7

2.a£3238E-09

O000or; 1 00000?

0002C,: ] 995:320

000511 C, 99167t

000706 0 987534

001018 0 983_15

001270 0 979315

001521 0 97523_

001771 0 971169

f

O

0

C

O

0

U

000000

995826

9916C9

987529

983_07

9_9302.,

97521_

97!143

1.000000

0.995570

0.991159

0.996768

0.992_97

0.9780_5

0.973712

0,969298

Mean ma:ntenance action title

1'4aint_,-,ance actions per hour

Maintenance time constraint

_.009900 0

0.99097_ 0

0,993976 0

0,991005 0

'6.98806! 0

0.985144 0

0.982253 Ci

0.97938_ 0

(t)

090000

001025

0020_3

003056

00&062

005062

006056

00_04Z,

(hr s}: 259.2

: - .95,_<z_E-0 _

!.000000

0 99_963

0 989957

0 984981

0 980035

o 975,19

'3 97023:

0 96_377

!

0

0

0

0

C,

0

C

009000

994961

989947

98_958

97999_

971,056

9701q2

965253

I

0

0

0

0

C

0

0

0@0000

993938

987914

9819 <r

975973

970057

96a177

958 _- DD3

ORIGINAL PAGE IS

OF POOR QUALITY
Mean maintenance action t_me (hr's)= 17725.8

Maintenance actions per' hour" : 5,6_I_9_E-05

Maintenance t_me constraint (t) = 8

i.000000 0.000000 1.000000

0.__99_0 _,n ,, 0.000000 O. 9q___880

C.999880 0.000000 0.999760

0.999820 0.000000 0.9996h0

_.999760 0.000000 0.B99520

0.999700 0.000000 0.999400

0.9996_0 0.OOC000 0.999281

0.399580 0.000000 0.999161

1.000000

0.999880

0,999760

0.999040

0.999E20

0.999/400

0.999281

0.999161

!.000000

0.999880

0.999760

0.999640

0.999520

0.999400

0.99£280

0.999100

_';-.£_1' ln_2 i t'Lt_iq_F;C= _C +. i©ii t i file

t,i_: i i_ t:ei',,__i';Ce act ]ons pe{" flour

tts:iiqtenat,ce time ,.:onstl'a[n +

: .000000 0 000:300

L:.:::TCI 9 OC:".-fl2C

• .996422 0 0028a6

_'+_._=sc""165 _r: 00&259

0,990929 0 005065

.988712 0 007065

3.-86517 0 000/4 J9

0.98/4341 0 0099/46

(hiLS}

• w

: 216.1

=_7487E, -Oq

8O

' , 300000

0 . 990?'

3 . 99- 6/48

0 . 99 C/+ 9,_

0 . 987355

0,98_+230

0,991120

0 .c:78023

!

C

0

0

0

O

0

0

@90900

795:_:

9938}9

99047_

987219

98917&

981039

977914

; 000000

O 9E5391

0 99080::

0 986236

0 981890

0 977165

0 972661

0 968177

H__r, , ._a] i-rtenance act

P!__._nt--2_i,_i-,ce actions

!:_{:2 ,t£i,Gr;,:-_ time con

[ . 000000

C

(I

O

0

C

0

996686

993405

990155

986937

983749

980593

977467

0

6

0

0

0

per hOL!F -

st:"aint ft',

0 000000

0 000412

0 000820

001226

001630

002030

002/428

00282_

t_rne ( l-;r s ) -:

.O0&

0

0

]

0

0

0

0

_Sn

O00000

993779

987589

981445

975341

989279

963256

95727_

0

0

6

0

0

0

OO{COO

99:772

937582

9,81/43_'

975321

9692_7

963211

957213

0

0

0

0

0

i 000003

0 993362

986768

980218

973712

9672_8

960828

95/4450

ORtG;NAL PAGE IS

OF POOR QUALITY