UNIVERSITYof

NOTRE DAME

NASA/USRA UNIVERSITY

ADVANCED DESIGN PROGRAM

1989-1990

UNIVERSITY SPONSOR

BOEING COMMERCIAL AIRPLANE COMPANY

FINAL DESIGN PROPOSAL

THE SCREEM-J4D

A Proposal in Response to a Low Reynolds Number

Station Keeping Mission

May 1990

Department of Aerospace and Mechanical Engineering

University of Notre Dame

Notre Dame, IN 46556

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FINAL DESIGN PROPOSAL

Presented By

AEROSPACE ENGINEERS

FOR THE

TWENTY-FIRST CENTURY

07 May 1990

University of Notre Dame

Department of Aerospace and Mechanical Engineering
O SCREEM-J4D

i

FOR THE

AEROSPACE ENGINEERS FOR THE TWENTY-FIRST CENTURY

Eric Fick

Michael Park

Ed Dieser

Jeff Janicik

Steve Valentine

Rob Calloway

Chris Sturgis

Project Manager

Structures

Stability & Control

Stability & Control

Aerodynamics

Propulsion

Propulsion
TABLE OF CONTENTS

EXECUTIVE SUMMARY ..................................................... vi

3-View Drawing ................................................... viii

Specifications Summary ....................................... ix

INTRODUCTION .................................................................... x

I. PRELIMINARY CONCEPT DESIGN

A. Request for Proposals ............................................ 1-1

B. Mission Evaluation ............................................... 1-4

C. Design Requirements and Objectives ............... 1-5

D. Concept Selection Study ....................................... 1-6

11. AERODYNAMIC CONFIGURATION

A. Airfol Selection ...................................................... 2-1

B. Wing Geometry ..................................................... 2-4

C. Drag Estimation .................................................... 2-9

RL STABILITY AND CONTROL

A. Horizontal Tail ....................................................... 3-1

B. Vertical Tail ............................................................. 3-6

C. Dihedral .................................................................... 3-9

IV. PROPULSION SYSTEM

A. Motor Selection ...................................................... 4-1

B. Propeller Data Generation .................................. 4-2

C. Battery / Propeller Selection ............................... 4-4

D. Speed Control ....................................................... 4-10

PRECEDING PAGE BLANK NOT FILMED

iv
Va STRUCTURES

A. Structural Configuration ..................................... 5-1

B. Spar Configuration ................................................ 5-2

C. Materials Selection ................................................ 5-4

D. V-n Diagram ........................................................... 5-4

VI. WEIGHTS ANALYSIS

A. Weights Estimation ............................................... 6-1

B. Internal Layout ....................................................... 6-2

C. Center of Gravity Estimation .............................. 6-5

VII. PERFORMANCE

A. Lift, Drag and Lift/Drag ........................................ 7-1

B. Take-off and Turning Flight ............................... 7-1

C. Endurance and Range ........................................... 7-2

D. Landing .................................................................... 7-3

VIII. TECHNOLOGY DEMONSTRATOR

A. Construction Phase ............................................... 8-1

B. Flight Tests .............................................................. 8-2

C. Instructions For Pilot ............................................ 8-3

D. Safety Considerations ........................................... 8-5

E. Cost Analysis .......................................................... 8-6

F. Critical Data Summary ......................................... 8-7

IX. CONCLUSIONS

A. Disadvantages and Advantages .......................... 9-1

B. Applications ............................................................ 9-1

C. Final Word From The Design Team ................. 9-2

REFERENCES

APPENDIX
EXECUTIVE SUMMARY

The Screem-J4Dis a remotely piloted airplane that has been designed to fly at

a chord Reynold's number of 100,000 while performing figure-8 maneuvers

in a restricted area. It has a high aspect ratio main wing with a conventional

empennage giving it a "sailplane" appearance. A specifications table and

three-view drawing accompany this summary.

The flight plan calls for ascent to cruise altitude at 20 ft and then perform

three figure-8 turns around pylons. These pylons will be separated by fifty

yards and stationed inside a sports facility. Once completed, the pilot is to

make use of any remaining power by loitering before landing the plane.

The propulsion system of the J4D consists of a propeller-electric motor

combination with the prop mounted at the front of the fuselage. The

Tornado 10 inch diameter, 6 inch pitch, two-bladed propeller is powered by an

Astro 05 electric engine with 7 AA Nickel-Cadmium batteries. The system is

capable of 50 watts of power and has throttling capabilities. Of the propellers

with available data, the 10-6 was best fitted for the take-off distance and

maximum current draw constraints. The 05 engine was chosen for being

most lightweight while still supplying adequate power.

In order to provide sufficient lift for low speed travel, the J4D has an aspect

ratio of 11.72 with an 8.2 inch mean chord. The wing consists of a spar and rib

construction with Micafilm skin. Its low-mount and dihedral in combination

with the vertical tail were prescribed to allow for excellent maneuverability.

A major problem, however, is that due to the nature of a low-speed mission,

there is little margin for error between the cruise and stall velocities. As far as

center of gravity travel is concerned, a square fuselage will internally contain

the servos, engine, and so that the CG of the airplane is kept at about 33% of

the chord.

A combination of directional and longitudinal control will enable the J4D to

perform the figure-8 maneuvers. However, in order to avoid the

vi
construction and servo weight of ailerons, the rudder was designed to be over

one-half the size of the vertical tail to insure that the proper roll motion

could be attained.

After a semester of hard work and discussion, the members of Group D feel

that this proposal is a well-established answer to the station keeping mission.

We are confident in the success of our design and believe that the Screem-J4D

is worthy of continued support.

Aerospace Engineers for the Twenty-First Century

vii
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SCREEM- J4D SPECIFICATIONS

Endurance •

Range

Stall Velocity

Maximum

Maximum

Maximum

Maxamum

Maximum

Minimum

Mimmum

Minimum

PERFORMANCE

Velocity

Rate of Climb

Power Available

Bank Angle

Roll Rate

Turning Radius

Glide Angle

Takeoff Distance

3.9 min @ 23 ft/s

5500 ft @ 23 ft/s

18.9 ft/s

58.6 ft/s

6.3 ft/s

50 Watts

30 degrees

28.8 degrees/s

28.6 ft

- 4.1 degrees

42 ft

Maximum C1

Maximum Angle of Attack

1.3

14.5 degrees

CONFIGURATION

Wing Span

RPV Length

Weight

Wing

Airfoil Section

Angle of Incidence

Aspect Ratio

Chord

Dihedral

Taper Ratio

Fuselage

Cross section

Payload Volume

Empennage Airfoil Sections

Horizontal Tail Area

Elevator Area

Vertical Tail Area

Rudder Area

Motor

Battery Pack

Propeller

96 in (8 ft)

37 in (3.1 ft)

3 lbs (48 oz)

NACA 4415

3.5 degrees

11.72

8.2 in

10 degrees

1.0

4in X4in

25.17 cu in

Flat Plate

.63 sq ft

.21 sq ft

.38 sq ft

.21 sq ft

Astro 05

7 AA Nickel-Cadmium

Tornado 10-6
INTRODUCTION

i i

Upon receiving the request for proposals on January 18, 1990, in regards to a

low Reynold's number station keeping flight vehicle, Design Group D began

their design phase in pursuit of such a vehicle. Each member presented his

ideas in regards to the mission requirements and goals which facilitated the

group in their formal presentation of the Design Requirements and

Objectives. The individuals also presented various conceptual designs from

which the group as a whole chose one and then discussed modifications.

Each member then concentrated his effortson a particular aspect of the design

phase. The group chose to divide responsibilities such that members

assumed roles in the technical areas of aerodynamics, stability and control,

propulsion, and structures. Each conducted extensive trade studies which

shaped the design of the aircraft. The group presented the status of their

design proposal in a comprehensive Preliminary Design Review on March

28,1990.

Upon receiving approval for their design, the group began the manufacturing

and formal documentation phases. The manufacturing began April 17, 1990,

and progressed rather rapidly due to the assigning of a chief engineer and the

eagerness of the other members. The aircraftassembly was complete on April

29 and was rolled out for ground testing on May I, 1990. The aircraftwas first

flown successfully on May 4, 1990. In regards to the documentation of the

design, each member contributed his expertise and studies to this final

written report. It is presented as an overview of this design process and as a

guide for other design teams to follow.

The members of Aerospace Design Group D have chosen to change their

name to one that is less generic and more suggesting of their direction. They

are now referred to as Aerospace Engineers for the Twenty-First Century,

which is abbreviated AETC. Likewise, they have designated their aircraft the

SCREEM-J4D. This is an acronym derived from the first letter in each

member's name: Steve, Chris, Rob, Eric, Ed, Mike, and Jeff. The "4D" stands

for "For Diploma". It is often referred to as the J4D.

PRECEDING PAGE BLANK NOT FILMED
This document represents the Final Design Proposal as presented by the

Aerospace Engineers for the Twenty-First Century in response to the request

for proposals for a low Reynold's number aircraft. It progresses in sections

with the first being a look at the preliminary concepts. The next five sections

deal with the particular technical areas related to the design - aerodynamics,

stability and control, propulsion, weights analysis, and structures.

Performance parameters are then considered followed by discussion related to

the technology demonstrator. Discussion of the design ends with the

conclusions section and a final word from the design team. The references

and appendices represent good sources of information for future design

teams.

xi
SECTION I

PRELIMINARY CONCEPT

DESIGN
A. Request for Proposals

The following Request for Proposals (RFP) provided the design requirements

for a remotely-piloted vehicle (RPV) to be created by AETC.

UNIVERSITY OF NOTRE DAME

DEPARTMENT OF AEROSPACE AND MECHANICAL ENGINEERING

AE441: Aerospace Design: Request for Proposals

Spring 1990

FLIGHT AT VERY LOW REYNOLDS

NUMBERS - A STATION KEEPING MISSION

OPPORTUNITY

Most conventional flight vehicles are designed to operate in a flight regime

such that the Reynolds number based on mean wing chord are in excess of 106

and some currently approaching 108. Recently there has been interest

expressed in vehicles which would operate at much lower Reynolds

numbers, less than 105 . Particular applications are low speed flight at very

high altitudes, low altitude flight of very small aircraft and flight in other

planets' atmospheres such as Mars. There are many unique problems

associated with low speed flight which pose challenges to the aircraft designer

and which must be addressed in order to understand how to exploit this low

Reynolds number regime. Since many of the anticipated missions for this

type of aircraft are unmanned, it is necessary to couple developments in

unmanned aircraft development with our knowledge of low Reynolds

number aerodynamics in order to develop an aircraft which can fly as slow as

possible at sea level conditions. This study will help to better understand the

problems associated with flight at these very low Reynolds numbers.

Considering the potential applications, the aircraft must be very robust in its

control and be highly durable.

Preliminary Concept Design 1 - 1
OBIE S

1. Develop a proposal for an aircraft and associated flight control system

which must be able to:

a. Maintain level controlled flight and fly a closed course at flight

speeds corresponding to Reynolds numbers less than 2x10 5 and as close to

lx10 5 as possible. The greatest measure of merit is associated with achieving

the lowest mean chord Reynolds number possible and maximizing the loiter

time on a dosed course.

b. Be maneuverable and controllable so that itcan fly a dosed pattern

and remain within a limited airspace.

c. Use a propulsion system which is non-airbreathing and does not

emit any mass (i.e.rocket, etc.).

d. Be able to be remotely controlled by a pilot with minimal flying

experience or an autonomous onboard control system.

e. Carry an instrument package payload which weighs 2.0 oz and is

2"x2"x2" in size.

2. Take full advantage of the latest technologies associated with lightweight,

low cost radio controlled aircraft and unconventional propulsion systems.

3. All possible considerations must be taken to avoid damage to

surroundings or personal injury in case of system malfunction.

4. Develop a flying prototype for the system defined above. The prototype

must be capable of demonstrating flight worthiness of the basic vehicle and

flight control system. The prototype will be required to fly a closed figure "8"

course within a highly constrained envelope. A basic test program for the

prototype must be developed and demonstrated with flight tests.

5. Evaluate the feasibility of the extension of the aircraft developed under

this project to high altitude station keeping application for atmospheric

sampling.

SYSTEM REQUIREMENTS AND CONSTRAINTS

The system designed shall satisfy the following.

a. All basic operation will be line-of-sight with a fixed ground based pilot,

although automatic control or other systems can be considered.

b. The aircraft must be able to take-off from the ground and land on the

ground.

Pmlimin_/Concept Design 1 - 2
c. The aircraftmust be able to maximize the loiter time within a restricted

altitude range on a figure "8" course with a spacing of 150 ft between two

pylons which define the course.

d. Ground handling and system operation must be able to be accomplished

by two people.

e. The complete aircraftmust be able to be disassembled for transportation

and storage and fitwithin a storage container no larger than 2'x2'x4'.

f. Safety considerations for systems operation are critical.A complete safety

assessment for the system is required.

SPECIAL CONSIDERATIONS FOR THE TECHNOLOGY DEMONSTRATOR

a. The Technology Demonstrator will be a full sized prototype of the actual

design.

b. The flight tests for the Technology Demonstrator will be conducted in the

Loftus Center on a closed course similar to the one described above. The

Demonstrator will be required to complete 3 laps on the course. The altitude

must not exceed 25' at any point on the course.

c. Takeoff must be accomplished within the 150' takeoff region shown on the

enclosed figure.

d. Loiter time will be based on the time needed to complete the 3 complete

laps in the air.

e. The design team must make provisions for estimating altitude and flight

speed during the tests. This information is to be monitored from ground

based observers.

f. The propulsion system for the Technology Demonstrator must not

contain any chemicals or any other substance which could prove harmful to

the Loftus Center or the aircraft operators.

g. The radio control system and the instrumentation package must be

removable and a complete system installation should be able to be

accomplished in 30 rnin.

h. System control for the flight demonstrator will be a Futaba 6FG radio

system with up to 4 $28 servos or a system of comparable weight and size.

i. All FAA and FCC regulations for operation of remotely piloted vehicles

must be complied with.

Preliminary Concept Design 1 - 3
B. Mission Evaluation

In order that the aircraft possess the ability to perform the basic mission, there

are several additional constraints regarding the aircraft and its mission which

need to be specified. In order to maintain a lightweight and relatively simple

airframe, no ailerons will be used. All banking and turning will thus be

accomplished by deflection of the rudder. Due to the possible complications

of flying into and out of ground effect, it is desirable for the aircraft to be

above the influence of ground effect before banking into the first turn.

Therefore, the aircraft must enter and leave all turns at the design cruise

altitude, losing or gaining as little altitude as possible. This may require

elevator deflections and/or throttle setting adjustments. For safety, the

aircraft will be flown at an altitude of 20 feet, which is 5 feet below the

absolute ceiling of the test area. This gives the pilot a margin (although

small) to make adjustments to save the plane in the event it would act

unpredictedly.

Combining these two constraints, it is evident that upon take-off the aircraft

should climb quickly to its cruise altitude of 20 feet and remain there for the

duration of the mission. By somewhat arbitrarily selecting the take-off

distance to be 75 feet (which is one-half of the available runway), it becomes

necessary for the aircraft to climb at a flight path angle of 15 degrees to reach

the cruise altitude of 20 feet over a horizontal distance of 75 feet.

As stated in the RFP, the flight envelope is very constrained, and as such the

aircraft needs to be very maneuverable. With that in mind, AETC feels that

the aircraft must have the capability to complete the turns using a 40 foot

radius, leaving no less than 20 feet between the aircraft and any side boundary

of the flight envelope during the duration of the mission. Since a large loiter

time is the greatest figure of merit, flying the mission at the lowest possible

velocity is a must. But everything is related, and since wing surface area (and

thus weight) go up as velocity goes down, a compromise exists somewhere in

between. It is our goal to find that compromise.

Preliminary Concept Design I - 4
C. Design Requirements and Objectives

In order to insure that the minimum goals defined in the RFP as well as

those specified in the above mission evaluation are satisfied, AETC has

imposed some additional (or more stringent) constraints on its RPV. They

are:

1. Maintain controllability while taxiing. Since the aircraft must be ground

launched, it is necessary for it to be steered while still on the ground at speeds

lower than those necessary for the lifting surfaces to provide this directional

control. For this reason, the landing gear must be coupled with the avionics

system to avoid the necessity of an additional servo to control the ground

steering.

2. Achieve airborne flight after a takeoff run of no more than 75 feet. This

constraint forces the aircraft to be off the ground and climbing with no less

than half of the runway left. This means that should some problem occur

immediately upon takeoff, there will still be sufficient runway left to put the

aircraft back on the ground without running it into a wall. Additionally, this

gives the aircraft a relatively long time to reach its cruise altitude of 20 feet;

mandating a flight path angle of approximately 15 degrees.

3. Climb at no less than 6 feet per second. Based on a cruise velocity of 23 feet

per second; if the aircraft takes off in 75 feet this rate of climb will allow the

aircraft to climb to its cruise altitude of 20 feet by the end of the runway.

4. Maintain controllable flight at a chord Reynolds number of no more than

lxl05. This is the lower limit of the Reynolds numbers targeted in the RFP.

Setting the design point at this Reynolds number ensures that should small

increases in velocity become necessary to avoid stall, the Reynolds number

will not be pushed over the 2x105 upper limit.

5. Maintain altitude during turns of no less than 270 degrees on no more

than a 40 foot radius at bank angles of up to 30 degrees. This constraint

evolved from safety considerations for the operator of the aircraft and for the

preservation of the aircraft itself. The figure "8" course mandates the turns of

at least 270 degrees; without these turns the aircraft will surely find its way

into a wall. The radius requirement ensures that the aircraft will possess the

ability to turn well within the maximum allowable radius of 60 feet given by

the geometry of the test flight area. Finally, the bank angle constraint ensures

Preliminary Concept Design | - 5
that even at severe bank angles, the aircraft will be in no danger of dragging a

wing on the ground due to lost altitude. Please note that in order to

max/mize the loiter time, the final test flight may not be flown at these target

values, but the aircraft must possess the ability to fly there should the

situation or need present itself.

6. Weigh no more than 3 pounds. This constraint evolved from

comparisons with other remotely piloted vehicles of similar size or mission.

The primary consideration here was that a lighter weight would allow for a

lower velocity and subsequently higher loiter time.

These constraints provided the guidelines by which AETC's aircraft, the

SCREEM-J4D, was designed and ultimately constructed. While not

specifically directed at any one area or sub-system of the aircraft, these

constraints had far-reaching implications in all aspects of the aircraft's design

and construction. Table 1-1 is a summary of these constraints.

TABLE 1-1: Design Constraints for the SCREEM-J4D

1. Maintain controllability while taxiing.

2. Take-off in distance <75 ft.

3. Climb at > 6 ft/s and fly at altitude = 20 ft.

4. Maintain contollable flight at a chord

Reynolds number < lx105.

5. Maintain altitude during turns >_..270 degrees

with turn radius < 40 ft at bank angles <30 degrees.

6. Weight <3 lbs.

D. Concept Selection Study

The selection of a concept became a laborious task for AETC once the Design

Requirements and Objectives (DR&O) were completed. The constraints given

by the RFP in addition to those in the DR&O had to be weighed very carefully

in the process of selecting a final concept. The approach that AETC took was

Preliminary Concept Design 1 - 6
considering three different concepts which were unrelated to each other but

were proposed by at least one member of the group. This was reasonable due

to the fact that each concept was considered in a more general form with

relation to the constraints. A more detailed analysis followed once one of the

three concepts was selected.

One of these concepts was a biplane configuration with a conventional

empennage (Figure 1-1). The main advantage of the biplane was that it could

provide extra lift which is vital to flying at low Re. Nonetheless, the

disadvantages were too numerous. The complications in resolving the

interference effects, the construction and connection of the two wings, and

the additional weight with the added wing were the main disadvantages

which prompted our decision to reject the biplane concept.

Another concept which AETC labored over for about a week was a canard

configuration with a vertical tail (Figure 1-2). The canard was considered able

to provide extra lift while maintaining maneuverability and not adding

weight. However, after doing further research on low speed RPV flight,

AETC discovered that the canard would have to be too large in order to have

a significant contribution toward the total lift. As a result, the airplane would

have just as much if not more wing surface area as a conventional airplane

without the convenience of the classical stability equations. Thus, the canard

concept was considered "pretty" but not the most efficient concept for the

mission.

A third concept which AETC decided to chose for the mission was the

"sailplane" type airplane with a conventional empennage. The most

important factor in the selection of this concept was that the theories and

equations which were such a big part of undergraduate aerospace study could

be most readily applied to the conventional airplane. In addition, the

learning process of the design and construction of this concept is useful in

gaining a better understanding of the process involved with the 747's, DC-10's,

etc., becoming a realistic part of the transportation industry.

Table 1-2 is a comparison matrix of the three concepts with regard to several

critical factors found in the DR&O or assumed from low speed flight. A +1

Preliminary Concept Desisn 1 - 7
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denotes that the design would benefit from the concept while a -1 denotes

that it would be detrimental. This table enabled AETC to be more easily

convinced that selecting a high aspect ratio conventional airplane as their

concept would best suit the mission.

TABLE 1-2 : Comparison Matrix for Concepts Considered

FACTOR BI-PLANE CANARD CONVENTIONAL

Added Lift +1 0 0

Ease of Construction -1 +1 +1

Weight -1 0 0

Use of Equations -1 -1 +1

Maneuverability - 1 - 1 - 1

TOTAL -3 -1 +1

The initial analysis for providing a little more detail to the concept was based

on the Reynold's number constraint of 100,000 and the weight constraint of 3

pounds. Given that the Re is proportional to the mean chord length and the

velocity and inversely proportional to the kinematic viscosity, it was seen

that by selecting a target Re and varying the mean chord, a velocity which

satisfied the constraint would be produced. Using this velocity and chord and

estimating a target lift coefficient, the span necessary to produce the required

lift was calculated. This process was aided by use of Computer Program A

which is located in Appendix I. However, knowing that no piece of the

aircraft can be longer that 4 feet required that the span be kept reasonably low

in order to avoid having a wing with too many sections. Using this method

with the lift coefficient of 0.9 and a weight of 3 pounds, it was found that the

allowed velocity and required span for a chord of 5 inches was 40 ft/s and 4.2

feet, respectively; while for a 12 inch chord, the same quantities would be 16

ft/s and 11 ft.

Because the loiter time is dependent on velocity, the first case is not optimal,

and because lightweight is of critical importance, the large wing area of the

Preliminary Concept Design 1 -8
second case, with its smaller velocity, is not optimal. Obviously, AETC's

solution fell between these cases.

AETC made seveal other decisions about the conventional concept once it

became obvious it was best suited for the mission. In addition to the high

aspect ratio wings, AETC decided against tapering the wing because the chord

at the tips would become too low and possibly introduce problems with

extremely low Re flight. Additionally, an untapered wing is much easier to

construct since the airfoil sections are all the same size. Ailerons were

discarded as a possibility because the added weight of a servo and the

construction required outweighed the advantages of better roll performance.

Past RPV's and research proved that the same roll performance could be

attained with just the elevator and rudder in the first place. Sweep is not

advantageous for low speed flight since it reduces the effective velocity the

wing sees. However, dihedral aids in roll control. Winglets were considered

in the concept study in order to reduce induced drag. But when the high

aspect ratio of the wing was considered, it was decided that the effect of

winglets would be too small to justify the added weight and construction

difficulties. A low-mount wing was selected to help destabilize the roll which

would facilitate the figure-8 maneuvers.

The fuselage was selected to be a box configuration for ease of construction

and containment of the propulsion system. The aft section of the fuselage

was tapered to reduce weight and skin friction drag.

The empennage was designed to contain two control surfaces -- rudder and

elevator. A relatively large rudder (55% of vertical tail area) was needed to

produce the yaw required to attain the desired roll moment. The rudder, in

conjunction with a large wing dihedral (10 degrees), enabled the aircraft to

perform the figure-8 turns as desired. The elevator comprised 33% of the

horizontal tail area, ensuring positive longitudinal stability and control. The

center of gravity was placed behind the aerodynamic center so that the

horizontal tail would produce lift rather than a downward force.

Lastly, the landing gear was comprised of three wheels - one on either side of

the fuselage mounted to the wing and one in the rear as a tail dragger. The

Preliminary Concept Design 1 - 9
two on the wing were spaced well enough to prevent tipping during take-off.

The tail dragger was connected to the rudder so that one servo could be used

for turning control during ground roll and flight.

These characteristics of the basic concept may be seen in the three view

drawing in the Executive Summary.

PreliminaryConcept Desisn 1 -I0
SECTION II

AERODYNAMIC

CONFIGURATION
A. Ai_ofl Selection

At low Reynolds numbers, viscous effects become more significant with

respect to inertial effects. The viscous region extends beyond the surface of

the body both in the lateral and longitudinal direction. The wake retards the

oncoming freestream, while interaction between the viscous and inviscid

regions is strong and nonlinear. Therefore, small changes in the position

(i.e., attack angle) of the body exposed to the flow will cause significant

changes in the surface pressure distribution. Due to the effects of this viscous

region, the flow is less capable of negotiating an increase in pressure. The

adverse pressure gradient can grow large enough to affect a peculiar

aerodynamic phenomenon--a separation "bubble," which is an enclosed

region of stagnant air that develops on the upper surface(s).

The flow over a smooth airfoil at low Reynolds numbers is often laminar,

unless means are employed to disturb the flow and thereby induce

turbulence. (Disturbing the flow energizes the boundary layer, thereby

increasing inertial effects and delaying separation). As the Reynolds number

decreases in laminar (undisturbed) flow, a separation bubble develops on the

upper surface and effectively increases the camber of the airfoil, thus

providing additional lift. However, the separation bubble is extremely

sensitive to disturbances (such as surface roughness and sound), and/or an

increase in the angle of attack. If the separation bubble breaks down, the

pressure drag will increase drastically. In addition, hysteresis is particularly

significant in airfoil data taken at low Reynolds numbers and at high angles

of attack. Therefore, the airfoil performance depends strongly on whether the

angle of attack is increasing or decreasing. Thus, in selecting an airfoil, it is

important to keep in mind the characteristics of that airfoil in the low

Reynolds number regime, as well as the flight plan dictated by the mission

requirements. In particular, selection criterion should include gradual stall,

high stall angle and/or low stall speed, and lift performance which suits the

flight plan and aircraft gross weight.

Measures of Merit: To select the airfoil intended for operation in a low

Reynolds number regime, the following characteristics are necessary:

1. Favorable lift curve slope and high CLmax

Aerodynamic Configuration 2 - 1
2. Gradual stall

3. High lift/drag ratio

4. Ease of fabrication

Design Constraints and Considerations: There are several fixed aerodynamic

performance parameters which are influential in the design of the airfoil.

The following were most critical to the J4D:

1. Cruise Reynolds number (based on mean wing chord), Rec=105

2. Cruise velocity, V_ = 23 ft/s

3. Mean wing chord, c = 0.683 in.

4. Freestream density, p=2.37 x 10 -3 slugs/ft 3

Besides these constraints, there are two are other important guidelines which

were used during the design phase. The first was the capability of precise

fabrication of the selected airfoil. For example, very thin airfoils, although

popular for low Reynolds number operation, are difficult to construct

precisely with the materials and tools which were available to AETC. In

addition, thin airfoil sections are subjected to higher stresses in flight and

under static conditions. The second important consideration was that stall

characteristics must be amenable to the desired flight conditions. Given the

requirements for lift performance, the maximum lift desired must be

achievable without flying at a geometric angle of attack close to the stall angle

((Zstall). Therefore, finite lift characteristics yielded by the airfoil parameters

and the wing geometry must be acceptable for the mission requirements

while restraining the wing size to reasonable limits to minimize the aircraft

overall weight.

Assumptions: Because the airfoil data used in the selection process [Ref. 12]

was taken at Reynolds numbers close to, but not exactly, the design operating

Reynolds number for our vehicle (105), interpolation was necessary to

determine the important parameters (zero-lift angle of attack, Clmax, lift curve

slope, IZstall, and maximum L/D). Figure 2-1 is an indication of the estimated

data for Re= 105, for the NACA 4415. A linear variation with Reynolds

number was assumed for determining this data. In addition, hysteresis effects

were neglected since the intended flight plan is a steady state cruise at a

constant 0_.

Aerodynamic Configuration 2 - 2
1--i

!

,,m',,

Z

i

i !i

m_t - l_Imlo!lJ,o2 'iue,.uo N

!

b

!

t

t

I

I

l

.v!

.v .j- _,

.'.% _i

""r'.. ".1\

t% %!

I. '_ll 41

¢1 O 0 O

II _I 'B

aol 'I;) - Ollltl h4111,1"l

tlt

e8

, , |

1

**'

i'"

_ I,

0

Q

Is

I

tl

i,<

i

O

U

i:

,4c

o,

0

'_ i.

';° i o.

Ill

0 ¢i !

qlP

I

e;3 - lilOl:!tlto;3 llx,i a

I I if i ." 3

u i::

IXXXX

m N_m_

0 I

wl

N

1 i/ """". _'_

" : ;I

i i

'i::) _ lUOlOllliO 2 fill

I

,t..,_"_

.-_

0

ciQ

,IP

N

I

t

I

:° 0

li

I

0

II

!

i'<

ll, i

o

li
Selection Procedure: To avoid the separation difficulties associated with

flight in the low Reynolds number regime, there are two primary design

options: (1) provide means for energizing the boundary layer and thereby

delaying separation, and (2) fly at geometric attack angles well below 0_$talI.

There are, however, disadvantages associated with each of these options.

Option (1) implies a substantial skin friction drag penalty, since the wall shear

stress is greater for turbulent flow. If option (2) is selected, the risk of

separation and stall are greater with an unexpected increase in 0_ (gust, nose

up). Because minimizing drag is an important objective in this design, and

because the desired lift performance can be achieved at angles of attack well

below _stall, option (2) was employed on the SCREEM-J4D.

The research for airfoils at low Reynolds numbers was conducted with

Reference 11, critically evaluating each airfoil based on the measures of merit

above. A number of laminar flow airfoils were rejected with relatively sharp

leading and trailing edges because of substantial fabrication difficulty and the

concern with high stresses in these thin regions. In addition, the precision

requirements of laminar flow airfoil construction are unrealistic given the

time and resources available to AETC. Very slight disturbances, particularly

at the leading edge, will produce undesirable effects if design option (2) is to be

used. (Option (1) is not compatible with laminar flow airfoils, due to the

associated skin friction drag penalty). Thus, although better lift characteristics

are produced in the low Reynolds number regime by laminar flow airfoils,

disturbance of the laminar flow is very difficult to avoid; therefore, a thicker

airfoil profile is desirable.

After a thorough evaluation of the low speed airfoils found in Reference 12,

two were selected for final analysis, the AH 79-100C and the NACA 4415 (see

Figures 2-2 - 2-8). A finite wing lifting-line analysis was performed for each

airfoil using the computer program listed as Item 2 in the Appendix, The

computer analysis requires the infinite lift curve slope and the zero-lift angle

of attack of each airfoil. These parameters, coupled with the wing planform

geometry, provide enough information to obtain reasonable predictions of

the lift and induced drag characteristics of the finite wing. It was found that

although the AH 79-100C achieved a better performance rating for the desired

Aerodynamic Configuration 2 - 3
Figure 2-2: Airfoils Considered

Scale Profile el U_e AH-75-100C Air_ot'l

0.5,i

0.25

O.IlO

4.;_S

.1-------_'--

CROSS SECTION OF NACA 4415 AIRFOIL

f

o.oo

,.o.2,,5

0+.

4,.5O

0.0 O.2 o_+ o_ o_ ,.o

x;c
¢D

o

cJ_

¢D

eq,,_

,-]Z! X

000

tl

<

rrx÷0U

o.

II Q

/

O0

0Q
!

L'_

,<

_f
A

8

0.018

0.016

0.014

0.012'

0.010

0.008

0.006

Figure 2-5. NACA 4415 CDi vs. Geometric Angle of Attack

J

(best-fit polynomial approximation)

_r

y = 2.4393e-

• • • • " • " " I " " " " I " " " " I • " ' " I •

2 3 4 5 6

Geometric AnsIe of Attack [de 8]

.J,a

.=.

_J

0.05

0.04

0.03

0.02

0.01

-1

Figure 2-6. NACA 4415 Drag Polar

y

I (best-fit polynomial approximation) l' I

1.7237e-2 - 1.6980e-2x + 9.0213e-2x^2 - 0.15822x^3 + 9.0225e-2x*4 R*2 = 0.987

inear region - constant lift sl

m____l__J _ll_rr_ion - sudden drag rise)

(C1 range) "_r

i

I |

0 1 2

Cl (infinite span)
==

(g

8

==

_5

u

0.016

0.014

0.012'

0.010

0.008

0.006

0.004

Figure 2-7. AH 79-100C CDi vs. Geometric Angle of Attack

0.002

1 7

./

(best-fit polynomial approximation)

/

y=2.O

I I l l l

2 3 4 5 6

Ceometri¢ AnKle of Attack [deg]

Figure 2-8. AH 79-100C Drag Polar

O.10 1 (best-fit polynomial approximation in operating range) !

==.

o.=

..=

_J

y = 5.2533e-2 - 5.9860e-2x + 0.23967x^2 - 0.40537x^3 + 0.24865x^4 - 4.4244e-2x^5 R^2 = 0.22

1

0.08

1 (range of lift coefficient values during operation)

0.06

0.04

0.02

0.0 1.0

C! (infinite span)
lift characteristics and the measures of merit above, it was not chosen as the

main wing airfoil. The fact that it is so thin introduced construction

difficulties and posed some concern for the airfoil's structural stability and

consistency. Consequently, the AH 79-100C's better performance

characteristics were not convincing enough to accept its disadvantages, and

the NACA 4415 was selected by default.

Final Selection: The NACA 4415, based on the information available in

Reference 12 and interpolated approximations for Re=105, has the following

approximate parameters at infinite span:

Lift curve slope: 3.72 per radian

Zero-lift angle of attack: - 4 degrees

Clrnax : ~ 1.2

L/Dmax : - 15.5

The values for maximum lift coefficient and maximum lift-to-drag ratio did,

of course, change when a finite aspect ratio was considered.

B. Wing Geometry

Important Parameters: To specify the wing geometry, the following

parameters must be determined:

1. taper

2. span

3. geometric and aerodynamic twist

4. dihedral (discussed in Section In)

These parameters were determined using these fixed (design) values:

1. mean-chord Reynolds number, Re= 105

2. cruise velocity, Vo.= 23 ft/s

3. freestream density, p,. 2.37 X 10 -3 slugs/ft 3

4. freestream viscosity, It= 3.76 X10-TN.s/m 2

The wing span and chord variation together determine the planform

[projected] wing area. The mean chord is fixed at a value of 0.68 ft by the

cruise velocity and desired Reynolds number.

Aerodynamic Configuration 2 - 4
Measures of Merit: The planform geometry was selected based on the

following characteristics:

1. finite lift curve slope

2. stall propagation pattern

3. spanwise lift distribution

4. induced drag

5. structural weight of wing and wing loading

It is desirable to obtain finite lift characteristics commensurate with the

performance requirements and overall system weight. To minimize the

overall system weight, the planform geometry was selected such that the

overall weight of the main wing was at a minimum, given the performance

requirements.

To obtain a favorable stall pattern, the planform geometry was selected such

that the maximum lift coefficient occurs at the root of the wing. The stall will

thus propagate from the root. This is particularly desirable for minimizing

bending moments in the wing. This stall pattern is also desirable if control

surfaces (i.e., ailerons) are to be located on the outboard wing sections.

Although the SCREEM-J4D was designed without these control surfaces,

minimizing wing bending moments is still extremely important.

An additional means of minimizing the stresses incurred in the wing is to

obtain a lift distribution which, when coupled with the wing surface area,

yields a reasonable wing loading. This total wing loading is obtained by

integrating the spanwise lift distribution.

Design Constraints and Considerations: As previously mentioned, airfoil

selection and wing geometry are influenced by several parameters which

were fixed by AETC in the interest of successfully meeting the mission

requirements. Table 2-1 is a summary of these fixed parameters.

Aerodynamic Configuration 2 - 5
TABLE 2-1: Design Parameters for Airfoil & Wing Geometry

PARAMETER VALUE

O_ 2.37 X I0 -3 slug/ft 3

V_ 23.0 ft/s

Rec 105

aircraft _ross weight 3.0 lbs ,

Other constraints and decisions were made during the preliminary design

concept (in addition to fixed parameters):

- stall angle - required CL must be achieved at 0_ < (Zstal 1

- no taper (taper ratio = 1.0)

no aerodynamic twist

If a taper ratio is employed on a wing intended for operation at low meanchord Reynolds numbers, separation effects will cause substantial lift

reduction and drag increase at the tip section. This is due to a lower chord

than the mean chord and hence the local Reynolds number is smaller than

the mean-chord Reynolds number. Therefore, the taper ratio of the wing will

remain fixed at 1.0.

Since the mean-chord is fixed by the Reynolds number and cruise velocity,

and the chord is constant along the entire span of the wing, the span is

determined by a choice of the planform area (or vice versa). Obviously, the

choice of airfoil section also directly affects the finite lift characteristics of the

wing. The [infinite] airfoil section parameters which affect the finite lift

characteristics are the lift curve slope and the zero-lift angle of attack. Because

the lift generated by the wing depends on its wetted surface area, the lift is

dependent on the choice of wing area or span. In this case, area and twist are

the only wing geometry parameters which need to be specified to generate a

certain lift force. Aerodynamic twist will not be employed, in the interest of

simple wing construction. Thus, our wing geometry options are limited to

span and geometric twist configuration. As a result, the independent

variables whose effects on lift have been investigated are the planform wing

Aerodynamic Configuration 2 - 6
area, geometric twist configuration, infinite lift curve slope, and infinite zerolift angle.

An important objective of AETC is to develop a design which is easily

constructed, yet capable of successfully meeting the mission requirements.

Thus, the independent variables investigated were chosen in part on the basis

of their relatively insignificant effects on construction capability. Because

aerodynamic twist complicates the wing construction, and (as it turns out) the

desired lift performance can be achieved without employing this technique, it

has been disregarded as a possibility in the design of the wing. Further, the

weight of the wing is an important consideration. Minimal weight of the

aircraft is particularly crucial in low-speed flight. Therefore, the wing size

must be sufficient to achieve the maximum desired lift force, yet no larger. It

is important to remember that because interference effects with the fuselage

are not taken into account in the lifting-line wing analysis, the performance

predictions of the actual wing will be slightly inferior to those calculated by

using finite wing theory.

Assumptions:

assumptions:

The analysis that follows is based on several important

1. The mean-chord Reynolds number is based on the cruise

velocity, which is fixed at 23 ft/s.

2. The angles of attack used in the analysis are sufficiently small for

valid applicability of the monoplane [finite wing] equation (Eqn.

2.58 [Ref. 2], Eqn. 5.51 [Ref. 1]).

3. Symmetric load distribution.

4. Effects of fuselage on wing lift performance are not taken into

account.

Methods of Analysis and Selection: To determine the span efficiency factor,

finite lift curve slope, induced drag vs. geometric angle of attack, and

spanwise lift distribution, a finite wing analysis program was employed at

geometric attack angles in the range, 1 degree < 0_ < 11 degrees. The finite

wing analysis program employs the monoplane equation of Prandtrs lifting

line theory [Ref. 2]. This equation is:

Aerodynamic Configuration 2 - 7
2b sin n_..__._O0

a(0) - _(c(0)) _ (An sin nO + nan sin 0 ) + (ZL--0

where 0_(0) = geometric angle of attack

c_I.=0 = airfoil zero-lift angle of attack

b = wing span

c(0) = functional form of the chord variation (in our case, c(q) is

constant)

An = Fourier coefficients from the proposed Fourier series

representation of the circulation distribution.

To reach a decision regarding the optimum combination of the planform

geometry parameters considered, the required lift during the take-off and

cruising phases of the flight were determined based on weight estimation,

desired rate of climb, and an estimate of the losses incurred due to fuselagewing interference. The desired CL for cruise was found according to:

W

CL = 1

p V,., 2 (c) (1)

where W is the total aircraft weight, and (c) is the mean chord (which is fixed

by the design). The (finite) lift coefficient required during take-off was also

computed with this equation replacing W with the required lift during takeoff. The required lift was determined from the following equation:

L = Wsin0,

where 0 is the desired flight trajectory. To account for the interference effects,

a factor of safety was taken to be 1.2 based on an estimate of the effective loss

of lift. The values for CL determined by the above equations were multiplied

by this factor to yield realistic requirements. Note that because the induced

drag coefficient, Cd,i, (and therefore the total wing drag coefficient) varies

quadratically with CL, it is desirable to obtain the minimum value of CL

necessary to provide the required takeoff performance and sustain the aircraft

Aerodynamic Configuration 2 - 8
in flight at the desired cruise speed. Thus, a compromise was made between

lift coefficient and induced drag coefficient.

Results: Given the measures of merit, the design constraints and the

performance characteristics of the NACA 4415 airfoil (see "Airfoil Selection"),

Table 2-2 represents the final wing geometric and aerodynamic parameters.

TABLE 2-2: Selected Wing Geometry

Chord, c= 0.683 ft (constant)

Span, b= 8.0 ft

Taper ratio, _.= 1.0

Aero. & Geo. twist:, none

Planform area, S= 5.46 ft2

Span efficiency factor, e=0.89

Average wing loading -. 0.55 lbf/ft 2

Finite lift curve slope, a= 0.073/deg

Analysis, of Figures: The finite lift curve slope for the NACA 4415 may be

estimated from Figure 2-3 as 0.0733/degree. It is evident, also, that the

maximum desired lift coefficient which is equal to 1.2 is achieved at an

incidence angle of -11.5 degrees. The desired cruise lift coefficient, 0.87, is

achieved at -7.5 degrees. As expected, the induced drag coefficient varies

parabolically with angle of attack. The spanwise lift distribution for the

NACA 4415 is shown in Figure 2-9 and Figure 2-10 for angles of attack ranging

from 1 to 10 degrees. As is evident from these plots, the distributions are

approximately elliptical.

C. Drag Estimation

Paramount in any aircraft's design phase is an accurate estimation of the drag

forces produced by the aircraft in flight. Since the power required and many

other parameters depend almost exclusively upon the aircraft's drag, this drag

estimation needs to be constantly updated as the aircraft nears the production

phase. The method used here is a combination of one presented in Mr.

Jensen's thesis "A Drag Prediction Methodology for low Reynolds Numbers

Flight Vehicles" [Ref. 6] and another presented by Dr. Nelson in Atmospheric

Flight Mechanics [Ref. 7].

Aerodynamic Configuration 2 - 9
MI

(j

Figure 2-9. NACA 4415 Finite Wing Lift Distribution

(AOA-- 1 to 5 degrees)

0.8 .......................................... 5deg. i

(1.6'

"'" 3deg ......... ", ". I

-'--2de 8 __ , ,

0.4' Ideg

0.2

(root) > (tip)

0.0

0.0 0.2 0.4 0.6 0.8 1.0

Y/(bl2) (non-dimensional spanwise position)

(AOA=geometric angle of attack)

.o.o.o.o.,

AOA=I deg

AOA=2 deg

AOA-._ deg

AOA=4 deg

AOA=5 deg

A

el

..=.

,,-1

r,j

Figure 2-10. NACA 4415 Finite Wing Lift Distribution

(AOA= 6 to 10 degrees)

ll° Ill fill" II I I "l,ll. I Ii i 111_ llll lll _I Ii _ I

............................ "10 deg

1.0 ...................... "--"9deg. "_. |

I ......... 8de_,.o ""-. "'. ]

0.81 -..,,,;., I

t

0.2

(root) ' " _ (tip

0.0/ • , • , • , • , • I

0.0 0.2 0.4 0,6 0.8 1.0

Yl(bl2) (non-dimensional spanwise location)

(AOA=geometric angle of attack)

.....-o.........

..........

AOA=6 dog

AOA=7 dog

AOA=8 dog

AOA=9 deg

AOA=I 0 dog
In general, the drag polar for an aircraft is made up of a parasite (or form) drag

component and a lift-induced drag component. The general formula for the

drag coefficient is then:

CD = CDo + CL2/e/tAR

where CDo and e are obtained from a method called the drag breakdown and

AR is the aspect ratio of the wing.

Determination of CDo: The value for the parasite drag coefficient, CDo, was

determined by summing contributions from all of the aircraft's components

and dividing the result by the reference area of the aircraft, in this case wing

planform area. In general, this quantity is then boosted by ten percent to

include interference effects. The general equation used to calculate CDo is

shown here:

CDo = _(CDn An)/Sref

The l'I's represent the individual components of the aircraft and the CD and A

of each FI are given by certain parameters pertaining to the shape and position

of the component. The drag breakdown for the aircraft is summarized in

Table 2-3.

TABLE 2-3 : Component Contribution to Parasite Drag Coefficient

Component CDn An (CDn An)

Wing .007 5.46 ft2 .03822

Fuselage .110 .111 ft 2 .01221

Vertical Tail .008 .38 ft 2 .00304

Horizontal Tail .008 .625 ft 2 .005

Landing Gear .017 5.46 ft 2 .09282

Sum (CDFI An) = .15129

From this summation, the preliminary CDo value (using the above equation)

comes to .0277, and after adding the 15% interference factor, it grows to .0319.

Aerodynamic Configuration 2 - 10
Determination of the Aircraft Efficiency Factor, e:

factor, e, can also obtained from a breakdown method.

following equation:

The aircraft efficiency

This method uses the

1/e = 1/ew + 1/eb + 1/eo

When using this method, ew and eb refer to the efficiency factors of the wing

and body, respectively, and eo is a "other" fudge factor to account for

interference effects. Using this method, the aircraft's efficiency factor comes

to .655. Another method introduced in Reference 6 is based on empirical data

and is a function of the aspect ratio only. The formula given for this estimate

is as follows:

e = 1.78(1 -.045(AR) .68) -.64

The value calculated from this empirical formula was .713. Since we have no

basis for judgement between these two factors and both come from reputable

sources, a decision was made to take the average of these two numbers and

use the efficiency factor thus gained. Our complete drag polar is then:

CD - .0319 + CL2/.682gAR

It is interesting to note here that in the calculation of the parasite drag

coefficient, the landing gear, because of its large reference area (the same as

the wing) and relatively high CD more than doubles the value of this

coefficient. Therefore, if some effort can be made to reduce the drag of the

landing gear the aircraft will benefit from a much-reduced drag coefficient.

Aerodynamic Configuration 2 - 11
SECTION III

STABILITY AND CONTROL
A. Horizontal Tail

The primary responsibility of the horizontal tail for this aircraft is to provide

longitudinal static stability while allowing the aircraft to be trimmed at a

relatively small positive angle of attack. A secondary consideration is that the

tail must not make the aircraft so statically stable as to be unresponsive to

control input. This section will discuss the process involved with the sizing

of the horizontal tail; figures of merit with regard to the primary

responsibility are the aircraft's trim angle of attack and total lift generated.

The figures of merit with regard to the second consideration are the stickfixed neutral point and the stick fixed static margin of the aircraft; both of

which are measures of the static stability of the aircraft.

Aircraft Moment Coefficients - Trim Angle : The computer program listed as

Item 3 in the Appendix was created to study the trends involved with the

longitudinal static stability of the aircraft and as such dealt primarily with the

moment coefficient of the aircraft about its center of gravity. The program

was used to compute the contributions of the wing, fuselage, and horizontal

tail to the aircraft's Cmo and Cma. Using the method found in Reference 8,

these components are calculated as shown in Table 3-1 (all quantities needed

for the following calculations are computed as needed within the program).

TABLE 3-1: Component Contributions to Cmo and Cma

Component

Wing

Fuselage

Tail

Cmo Contribution

Cmow = Cmacw + CUw(Xcg- Xac)/C

Cmof = (k2-kl)/36.5ScY.wf2(0tow+if)Ax

Cmot = TIVHCLat(r-o + iw - it)

I

Cma Contribution

Cm_w = CLaw(Xcg - Xac)/C

Crno. f = 1/363ScY_wf2(deu/da)Ax

Croat = -_VHCLctt(1 - de/da)

After finding the contributions due to all of these components, the Cmo and

Cma contributions were summed resulting in the equation for the moment

coefficient about the center of gravity of. the aircraft:

Cmcg = Cmo + Cma 0_

Stability and Control 3 - 1
Setting this equation equal to zero will give the angle of attack at which the

aircraft is "trimmed":

a = - Cmo/Cma

Trends & Sensitivities : It should be noted here that the J4D's fuselage

contributions to the trim angle of attack are constant and almost negligible,

and although the wing's contribution to the moment coefficient is certainly

not negligible, it can be treated as a constant for the purpose of the program.

By assuming that these values do not change significantly as horizontal tail

parameters change, the program was used to examine the tail's influence on

the previously mentioned factors. As can be readily seen from the boxed

equations above, the contributions due to the horizontal tail factor

considerably into the final output of this portion of the program. Specifically,

the volume ratio of the tail and the tail incidence angle seem to have a great

impact on the tail's contribution to the moment coefficients. At this point in

the design phase; structural, weight, and overall size requirements placed the

aerodynamic center of the horizontal tail approximately 2 feet behind the

center of gravity of the aircraft, which is located at 33% of the mean chord of

the wing. This effectively froze the length to the horizontal tail and made the

volume ratio of the tail a function of its surface area only. Thus, all of the

parameters which were previously strong functions of the volume ratio now

become strong functions of the tail surface area.

Once the accuracy of the computer program was checked with a series of hand

calculations, it was modified to calculate the trim angles for a number of

different tail surface areas and incidence angles. These values generated with

this program were then plotted to show the variation of trim angle as a

function of different surface areas and tail incidence angles. As is

demonstrated in Figure 3-1, the trim angle of attack increases with increasing

tail area and decreases with increasing tail incidence angle. The multiple

lines on this plot and the ones that follow are for different values of the tail

incidence angle. The trim angles shown on the graph are for the fuselage

reference line of the aircraft. The wing, due to its incidence angle, will see an

effective angle of attack 4.5 degrees higher than this value. The airfoil

selected for the wing stalls at approximately 14.5 degrees, so the highest

possible angle of attack for the aircraft is 10 degrees. The target trim angle of

Stability and Control 3 - 2
Figure 3-1

<Z

5

TRIM AMGLE OF ATTACK [des] us. TAIL AREA [_t_a]

m

TAIL AREA [f't_a]
attack for the aircraft was roughly 6 degrees, allowing close to 4 degrees for inflight maneuvering before encountering problems with stalling the aircraft.

Horizontal Tail Influence on Lift : The trim angle of attack, though, only

tells half of the story. Using this angle of attack as well as the downwash and

incidence angles, the lift due to the wing and tail were calculated and

summed. Figure 3-2 shows how the tall area and incidence angle influence

the lift characteristics of the aircraft. The basic trend is the same: the lift

increases with increasing tail area and decreases with increasing tail incidence

angle. Since this analysis is dealing only with the the "steady-level" flight

phase, the only condition that had to be met here is that the total lift must be

equal to 3 pounds. Any excess lift is not needed at this point, and any less lift

will not produce a flyable aircraft.

Stick Fixed Neutral Point & Static Margin : The second portion of this

analysis deals with the stick fixed neutral point and static margin of the

aircraft. The equations for these quantifies are given in Reference 8 as:

Stick Fixed Neutral Point:

XNp/c = Xac/C - Cmcd/CLaw + TIVHCLat/CL_w(I - d_/d(z)

Stick Fixed Static Margin:

SFSM-- X /c- Xcg/C

Although it is difficult to see from the equations standing alone, the variable

factor having the most influence on these quantities is the horizontal tail

volume ratio. Knowing that the tail length is effectively frozen, the size of

the horizontal tail is highlighted as the "most influential factor" in stability

considerations.

As a basis for comparison, the horizontal tail surface area was initially sized at

25% of the main lifting surface area (1.367 ft 2) and located 2 ft behind the C.G.

This combination created a statically stable aircraft capable of producing the

required lift at a 5.5 degree angle of attack. This configuration, however, also

produced a Cm¢z value of -2.6 per radian, along with a neutral point at 100%

of the mean chord (the end of the wing), giving a static margin of 67% of the

Stability and Control 3- 3
Figure 3-2

4

TOTAL LIFT GEHERATED [lb] us. TAIL AREA [_t_2]

i t

f'l

J_

L*

uJ

pn,

hi

Z:

UJ

kh

/

.J

<Z

p0

k0.4 0.6 0.8 1.0 1.2 1.4

-1.5

-i .0

-0.5

0.0

0.5

1.0

TAI L AREA [ F tI*2 ]
mean chord. While this would certainly be a stable aircraft, it would probably

prove to be very sluggish. In order to make the aircraft more responsive, it is

necessary to make the value of Cma less negative and to move the neutral

point closer to the center of gravity, thus effectively reducing the static

margin. Reverting back to the boxed equations, it is easy to see that by

decreasing the volume ratio (through the tail surface area) Cma will become

less and less negative. Figure 3-3 shows the variation of Cma with the tail

surface area while Figures 3-4 and 3-5 show how the neutral point in

percentage of mean chord and static margin are influenced by the same

quantity.

Results: With all of these observations in mind, the time for decision

approaches. The basic criterion remains the same: produce enough lift at a

relatively low angle of attack. But there was a new twist in the form of the

Cmc_, neutral point, and static margin values which are actual measures of

the aircraft's longitudinal static stability. Certainly, Cmct must be negative to

ensure static stability for the aircraft under consideration. Guidelines from

Reference 8 indicate that static margins of 5% of the mean chord are

acceptable targets for civilian aircraft. Our preliminary estimate was more

than thirteen times that value. More constraints were needed to reduce the

number of choices. Further investigation revealed that the drag increases

with increasing tail area, and that the lift-to-drag ratio is relatively

independent of the tail surface area once it rises above 0.8 square feet for all

tail incidence angles (see Figures 3-6 and 3-7). The overriding factor then

came from the structures side of the design team, who stated that it would be

much easier to mount the horizontal tail if the incidence angle were small

than if it were large. With this information in hand, the decisions were

made, and the point highlighted on the plots indicate the status of the final

design. Table 3-2 is a summary of these characteristics of the J4D:

Stability and Control 3 - 4
Figure 3-3

I,,11

t3

0.4

Cm)al_a vs. TAIL AREA [ft_2]

0.6

, I , I , I

0.8 I .0 I .2 1.4

TAIL AREA [_tZ_2]
Figure 3-4

NEUTRAL POINT us. TAIL AREA [_t_2]

1' I

8. l

8.a

8,4

, I , I , I

8.8 1.8 1 .::' 1.4

TAIL AREA [f t_Z2]
Figure 3-5

STATIC MC_QIH vs. TAIL AREA [_t_23

0.4 0.6 6,8

, I , I

1.0 1.2

TAIL AREA [Ft_23
Figure 3-6

TOTAL DRAG GEHERATED [lb] vs. TAIL AREA [_tZZ2]

0.30

i t

-I .5

-i .0

-0.5

0.0

0.5

1.0

0.10

0.4 0.6 0.8 1.0 1.2 1.4

TAIL AREA [_tZZ2]
Figure 3-7

15.0

12.5

LIFT TO DRAG RATIO vs. TAIL AREA [ftX*2]

0

p-q

0

I-4

.J

10.0

TAIL AREA [_t_2]
TABLE 3-2: Horizontal Tail Sizing Information

Volume Ratio .356

Length to Aero Center 2.005 ft

Planform Area .625 ft2

Chord .479 ft

Span 1.304 ft

Incidence angle -1 °

Cma - .685 rad -1

XNp/C .514

Static Margin .184

Aircraft Lift 3.0 It>

Aircraft 0_ 5.87 °

As is seen from this table, the static margin is reduced to 0.184, or just over 3

times the target value of 0.05 set in Reference 8. Additionally, the value of

C ma has been increased to allow for more control response and less

sluggishness. The aircraft produces the required lift and is trimmed at an

angle of attack of 5.87 degrees, allowing just over four degrees for

maneuvering before stall becomes a problem. Additionally, this tail surface

area places the aircraft in the region of relatively low drag and high lift to drag

ratio.

Elevator : The elevator consists of roughly one third of the horizontal tail's

surface area. Its surface area is .208 ft 2 and Cmse is an acceptable -0.683 rad -1.

This elevator area was initially sized based on the rules of thumb found in

Reference 5. Once the value of the elevator effectiveness was found to be

"large" enough, this value was accepted and finalized for the design.

Dynamic Stability : Under its final configuration, the aircraft is dynamically

stable as well; a damping ratio of .88 ensures steady flight and a time to halfamplitude of .105 seconds limits the time the aircraft spends recovering from

disturbances. The period of oscillatiori for these oscillations is 1.86 seconds,

meaning that the motion will damp out after less than one oscillation.

Clearly this aircraft will be able to perform its station keeping mission in a

Stability and Control 3 - 5
stable manner using equation found in Reference 8 modeling the aircraft as a

second order differential equation with constant coefficients.

B. Vertical Tail

Lateral stability and control are important considerations in the design of all

aircraft. For this rudder and elevator controlled RPV, the vertical tail and

wing dihedral were crucial elements influencing its stability, performance

and handling. Since the SCREEM-J4D was designed without ailerons, the

rudder was required to induce a sufficient yaw angle, 13, such that the vertical

tail and dihedral would achieve the roll maneuverability usually attributed to

the ailerons.

Initial Sizing Analysis: The primary consideration in the sizing of the

vertical tail and rudder was the ability to induce a side slip angle, or yaw

angle, [3 (beta). This induced yaw angle, coupled with the design wing

dihedral, will provide the lateral control for the J4D. The tail surface area

(Sv), flap effectiveness (_), and the maximum rudder deflection angle (Sr)

were initially selected comparable to typical RPV values. The initial Sv was

established through the combination of two "rules of thumb:"

Sv / Sh = 0.33- 0.40

Sv / Sw = 0.07 - 0.01

However, values of Sv outside the range of these typical values were also

considered due to the unusual objective of flying at low Reynolds numbers.

Table 3-3 contains a further listing of fixed and variable parameters, and the

typical RPV values used to select the initial values.

Stability and Control 3 -6
TABLE 3-3: Vertical Tail Sizing Information

Parameter

Surface Area (Sv)

Position aft CG (Iv)

Rudder Deflection (Sr)

Flap Effectiveness (z)

Tail Efficiency (nv)

Lift Curve Slope (Clav)

Value Reason

0..34- 0.60 ft2

1.5 - 2.0 ft

<25 °

0.62 - 0.75

1.0

2x

RPV typical values

required by horizontal tail

stall and structure stability

tail stability, rule of thuml

low wing in propwash

elliptical wing assumption

Lateral Moment Coefficient: The wing-body contribution to Cn_ was

calculated to be -7.21 X 10-5/deg through an empirical equation and is constant

throughout the analysis (Eqn. 2.74, [Ref. 8]). This destabilizing contribution

was negligible compared to the contribution of the vertical tail:

Cn_3v = Vv nv CLay (1 + da/d_) (Eqn. 2.80, [Ref. 8])

The value of nv(l+ dG/dB) was estimated to be 1.135 and directly

proportional to Sv (Eqn. 2.81, [Ref. 8]). In addition, since the vertical tail is a

fiat plate airfoil, the value of the lift curve slope was assumed to equal 2/1:.

The resulting Cn_ due to the vertical tail was calculated to be 8.4 X 10 -4

/degree, which is roughly ten times the magnitude of Cn_ due to the wingfuselage combination. The total yawing moment coefficient was computed

to equal 7.68 X 10 .4/deg through:

CnB = CnBv + Cnwf

The rudder control power, CnSr, was then calculated:

CnSr = -Vv nv z CLav (Eqn. 2.86, 2.87, [Ref. 8])

The ratio of dynamic pressures of the wing and the vertical tail, nv, was

assumed to remain constant at 1.0, due to the low-wing setting and the

positive effects of the propwash. From Figure 2.20 [Ref. 8], the flap

effectiveness z, of 0.72 was selected according the desired ratio of the control

surface area to the lifting surface area of 0.55. The rudder was sized at 55% of

the overall vertical tail to afford sufficient control power, without

Stability and Control 3 - 7
endangering structural stability. The rudder control power, CnSr, was thus

found to equal -5.33 X 10-4 /degree, at a maximum deflection angle of 25

degrees. From these parameter ranges, the overall yawing coefficient was

then determined to vary according to:

Cn = Cnl3 _ + CnSr 5r

In order to solve for the yaw angle, 13, at the trim condition, Cn was set equal

to zero. Including the trim condition constraint, Beta becomes directly

proportional to the rudder deflection angle; therefore, the optimum [3 occurs

with maximum 8r.

This relation verified and the location of the vertical tail aft of the CG,

selected in conjunction with the horizontal tail, the vertical tail surface area

(Sv) was swept through numerous values using the computer program listed

as Item 4 in the Appendix to achieve a satisfactory design. The vertical tail,

with tail position lv=1.9 ft and surface area Sv=.38 ft 2 (which includes a rudder

area of .21 ft 2), achieved sufficient yawing power to ensure safe completion of

the mission constraints while remaining in agreement with typical RPVs.

Numerous values of lv were analyzed and found to have no significant effect

on _. Since it's maximum effect is about +/- 0.1 degrees, it did not take any

precedence in the sizing of the vertical tail. The above parameters produced a

maximum yaw angle of 17.3 degrees at the maximum rudder deflection angle

of 25 degrees.

Roll vs. Yaw: Prior to finalizing the design to the above parameters, an

analysis relating the yaw to the roll was completed. The computer program

listed as Item 5 in the Appendix was compiled to compute the roll moment

due to yaw for various values of dihedral. This code relates the effects of

wing tip contour and aspect ratio to the roll moment coefficient through:

CI[_ = (CI_/I") I" + ACI_ ( Table 3.6, [Ref. 8])

Since the maximum ordinates of the wing will lie on the mean line, ACII3 is

assumed to equal zero. The roll moment coefficient due to yaw angle per

Stability and Control 3 - 8
degree, (Clj3/F), was extrapolated from Figure 3.9 in Reference 8, to equal -

.00027. Sweeping the dihedral angles from 4 to 12 degrees, over a range of yaw

angles of 0 to 30 degrees, a minimum dihedral angle of 7.5 degrees is required,

as shown in Figure 3-8. Note, the range of required roll moment coefficients

was initially determined by assuming ailerons. The ailerons assumed in the

design were selected using "rules of thumb" and RPV's of comparable size.

Thus, an estimate of the minimum required C1 of .035 was computed using:

C1 = (2 CLc_w t dod/Sb _ cy dy (Eqn. 2.96, [Ref. 8])

Before ending this analysis, it is important to note that although a decrease in

tail surface area shows an increase in [3, the maximum increase is

approximately +/- 1.0 degree. Still, a smaller surface area does denote less

weight and less skin friction drag. Thus an optimum design would include a

high percentage rudder of a smaller tail area (high "c, low Sv). Table 3-4

summarizes the vertical tail as designed for the J4D.

TABLE 3-4: Vertical Tail Summary

I

Surface Area 0.38 ft2

Aspect Ratio 1.38

Height 8.0 in

Root Chord 7.2 in

Tip Chord 3.9 in

Taper Ratio 0.54

Sweep 23.2 °

Aft CG Position 1.9 ft

C__ 7.68 x 10-4

Rudder Area 0.21 ft2

Max Rudder Deflection 25°

Flap Effectiveness 0.72

Cnsr -533 x 10-4

I

C. Dihedral

Four wing dihedral configurations were considered for the J4D. These

included a V-shape dihedral and three 3-panel polyhedrals, with various

Stability and Control 3- 9
Figure 3-8

C1 vs. BETA

0.000

Minimum Required Dihedral

-0 .Ba5 /

-8.873 ..

E

d

o -0.100

J, I I, I I

-0.1_'5 i i i i J , , , , , I , J i i

0 3 10 13 20 23 30

BETA (cle_ees)
panel break locations, as indicated in Table 3-5 and Figure 3-9. Functionally,

the dihedral links the yaw and roll axes of the aircraft, allowing a yaw

deflection to cause a rolling moment. This result is due to the increase of

angle of attack and thus an increase of lift on the forward wing. A comparable

decrease occurs on the aft wing. The net affect creates a wing generated

rolling moment. Note however, that the change of angle of attack of the

forward wing is equal and opposite to that of the aft wing. This equality

provided the wings with a constant total lift within the unstalled region.

TABLE 3-5: Wing Dihedrals

Dihedral T_)e

V-shape

3 panel

Inboard An_le Outboard An_le Panel Break

10.0 deg. 10.0 deg. N A

0.0 deg. 11.0 deg. 25% (1 ft)

0.0 deg. 11.9 deg. 37.5% (1.5 ft)

0.0 de_. I4.3 de_. 50% (2 ft)

Analysis and Results : The change in angle of attack (Aa), for the wing is a

function of the wing dihedral angle and the yaw input angle:

A(Z = tan -1 (SIN Yaw x TAN Dihedral) ([Ref. 4])

The results of the parametric sweep of the yaw angle from 1 to 25 degrees and

the dihedral angle from 7 to 21 degrees are plotted in Figure 3-10 (see the

computer program listed as Item 6 in the Appendix). The change in angle of

attack was plotted verses yaw angle for various Equivalent Dihedral Angles

(EDA, discussed below). As demonstrated in Figure 3-10, Act is proportional

to both the dihedral angle and the yaw angle. The performance envelope and

the design EDA are indicated. The envelop and the selected Aa are

constrained by the wing's stall angle, the maximum yaw angle of 17.3 degrees

and the minimum required dihedral angle of 7.5 degrees (above analysis).

The required wing angle of incidence for steady-state cruise is 10 degrees. The

stall angle for the wing at a Reynolds number of 100,000 is 14.5 degrees,

allowing a maximum Aa of 4.5 degrees. Due to the imperfections incurred

during manufacture, a maximum Aa of 4.0 degrees was selected.

Stability and Control 3 - 10
o£hed_ aZ

F£gu_ e

3-9

o

conf igu_at£°ns

couside_ed
Figure 3-10

O

O

O

w

.J

.J

W

RMGLE OF RTTRCK vs. YRW RI_ _or various WIMG DIHEDRRL [de_]

10

8

6

4

Flight Envelope

Gamma

21

2O

9

118

7

6

5

4

13

12

11

10

9

8

7

2

ign

0

0 5 10 15 20 a5

Y_d4 I_IGLE [de_rees)
Given the yaw angle of 17.3 degrees, corresponding to the maximum rudder

deflection of 25 degrees, an EDA of 10 degrees was selected from Figure 3-10.

EDA's of up to 15 degrees were also considered for this design, since they too,

satisfied the envelope constraints. However, the larger EDA's, while slightly

increasing the rolling rate, would decrease the wing's lift efficiency slightly

and increase the structural instability. The EDA of 10 degrees was also

ascetically more pleasing, thus increasing the J4D's marketability. While it

would be optimum to reduce the yaw angle, reducing the drag on the vertical

stabilizer during maneuvers, with a greater design EDA, 10 degrees was

selected to ensure an unstalled wing. Additionally, while the performance

requirements necessary for this mission do not "push" the limits of the

performance envelop for an EDA of 10 degrees; the V-shape dihedral

configuration was selected due to the minimum outboard angle, to reduce the

chance of tip stall, should the performance requirements be increased.

Roll Moments and Roll Rates : The maximum &0_ is shown to be 3 degrees

at the maximum yaw angle of 17.3 degrees for a 10 degree EDA. The change

in lift over the individual wing was computed, based on a constant lift curve

slope of 0.1 /deg within the wing's unstalled region. The total moment was

then summed from the section rolling moments computed from the change

in lift over 12 panels for an individual wing. Note, due to this dependence

on the change in lift, the rolling moment of the dihedral wing is

demonstrated to be independent of the overall angle of attack. A parametric

sweep of the rolling moment verses the total angle of attack for the leading

wing was computed (see the computer program listed as Item 7 in the

Appendix). The roll moment generated by the maximum AO_ of 3.2 degrees

is shown to equal 44.8 ft lb.

The rolling moment was then non-dimensionalized by computing the

moment fractions at their semi-span locations. Holding the outboard angle

(ODA), constant at 10 degrees, for reasons stated above, the rolling moment

percentages achievable at a maximum Ao_ were computed for the various

configurations (see Table 3-6). These rolling moment percentages correspond

directly to the EDA's of the various configurations according to Reference 4:

Stability and Control 3 - 11
Rolling Moment (%) x Outboard Dihedral Angle (degrees) = EDA (degrees)

TABLE 3-6: EDA vs. Roll Moment

Dihedral Type

V-shape

3 panel

Moment Fraction EDA Roll Moment

1.0 10.0 deg 44.8 ft lb

.91 9.1 deg 40.8 ft lb

.84 8.4 deg 37.6 ft lb

.70 7.0 de_ 31.4 ft lb

Note, the ODA is equal to the EDA for a V-shaped dihedral, since the panel

break occurs at centerline. The rolling moment for the V-shaped dihedral is

produced over the entire wing, while the fiat center panel of the 3-panel

configurations makes no contribution to the roll power, tasking the outboard

angles to generate the lift required to roll. Thus, for a constrained wing

loading and outboard dihedral angle, the V-shape dihedral, with the largest

EDA, will allow the greatest roll power. This is further demonstrated in Table

3-7, where given a constant yaw angle, the wing with the greatest EDA

develops the greatest roll rate.

TABLE 3-7: EDA vs. Yaw Angle

Yaw Angle 17.3 ° EDA 10 °

EDA Roll Rate Yaw Angle Roll Rate

10 deg 28.8 deg/s 5 deg 7.4 deg/s

9.1 deg 26.2 deg/s 10 deg 16.2 deg/s

8.4 deg 24.2 deg/s 15 deg 25.0 deg/s

7.0 deg 20.2 de_/s 20 de_ 31.7 de_/s

The roll rate performance was further analyzed with respect to yaw angle for

the design wing dihedral of 10 degrees. The data from this analysis is also

included in Table 3-7.

Stability and Control 3- 12
Turning Radius : This coupled relationship was further considered for

steady-state turning. Constrained to a 40 ft turning radius at the maximum

bank angle of 30 degrees, the aircraft was found to require a yaw angle of 5

degrees. Note, this is the required yaw angle to maintain a steady-state turn at

a turning radius of 40 ft. By increasing the yaw angle to its maximum of 17.3

degrees, the turning radius of the J4D can be minimized to 28.6 ft.

Stability and Control 3 - 13
SECTION IV

PROPULSION SYSTEM
A. Motor Selection

The first task in sizing a propulsion system was to select a motor. Since

design constraints prohibited the emission of particles into the atmosphere,

liquid-fueled engines of the piston or ducted-fan type were eliminated as

possibilities. Attention then turned to a propulsion system featuring a

propeller powered by an electric motor, which met the emissions constraint.

The task then was to select an electric motor.

The consideration of electric motors involved the Astro 035, Astro 05, Astro

15, and several generic "others". By generic "others" it is meant those motors

for which only the maximum power output, weight, size and RPM range

were known. No information regarding their operating characteristics were

available. These motors were typically discovered in modelling magazines or

manufacturers' literature. The motor selection process to be discussed is

summarized in Table 4-1. The first criterion upon which the selection was

based was power production. From initial estimates of motor power

requirements it was realized that the Astro 035 was inadequate. Its power

production was too low and hence eliminated from consideration. The

generic "other" engines all appeared attractive in terms of producing enough

power to meet requirements. However, since no information as to their

operating characteristics was available, any analysis and performance

predictions with them would be extremely difficult. This problem could have

been resolved through inquiries with the respective manufacturers or tests in

a laboratory. Time invested in either of these practices could have been costly

and possibly unrewarding. Due to time and practicality considerations, those

motors for which the operating characteristics were unavailable were

eliminated from consideration.

Attention then focused on the Astro 05 and Astro 15, whose operating

characteristics were known and available in tabular form. This allowed for a

performance analysis once a motor had been selected. In regards to the

selection, both produced an adequate amount of power; in fact, well above

that which initial estimates required. The size of both motors was attractive.

That is, the dimensions were such that both were compatible with the size of

the fuselage and would mount easily to the front of it. The deciding factor

Propulsion System 4 - 1
became the motor weight. The Astro 05 weighed an ounce less and its

standard (recommended) battery pack weighed 9 ounces less. Using smaller

battery packs than those recommended because of lower power requirements

would lessen the system weight for both. However, the Astro 05 would still

weigh less and therefore it was selected as the motor to power the aircraft.

TABLE 4-1 : Motor Selection Considerations

Astro 035 Astro 05 Astro 15 "Others"

Power Production : low adequate adequate adequate

NOT NOT

Specifications : available available available available

Size: good good good

Weight :

(Motor/System) 4.5 / 11.0 oz 65 / 16.0 oz 7.5 / 25.0 oz

B. Propeller Data Generation

In order to properly analyze the performance characteristics of different

propellers, it was found necessary to generate propeller performance

characteristics (thrust and power coefficients, Ct, Cp) as a function of advance

ratio J. Since performance estimates on hobby propellers were not available

and time did not permit actual wind tunnel testing, a data base was created

using a program entitled "Propeller Performance Analysis for Small

Computers," written by Barry N. Young [Ref. 11]. This code enabled one to

generate theoretical propeller performance characteristics based upon simple

blade element theory. Input to the program were airfoil data (of the propeller

blade section), estimated flight conditions, and the propeller geometry. These

geometrical quantities were the blade angle, chord, and thickness measured at

nine radial positions.

Propulsion System 4 - 2
Due to the lack of propeller data, it was assumed that all of the propellers used

a CLARK Y airfoil section which is used for the TopFlight family of hobby

propellers. It was also assumed that all of the propellers studied had a

constant blade angle distribution. Which dictated that the only geometrical

difference between propellers of the same diameter but different pitch was the

blade angle at each radial position. Thus, for a family of propellers with the

same diameter, the measurements only had to be performed once, with the

blade angle (13) determined by the pitch (P) according to the relation

P

= tan-1 2_r

where r is the radial position measured from the hub (P and r have the same

length units).

The propellers studied ranged in diameter (8, 9, 10, 11, 12, and 13 inches) and

pitch (4, 6, and 8 inches). Each of these diameters were studied with all three

pitches, hence, the total number of diameter-pitch combinations resulted in

the analysis of 18 different propellers. A sample of the input geometric

characteristics and parameters of Barry Young's program, in addition to a

sample of the theoretical performance results are listed as Item 8 in the

Appendix.

Simple blade element theory (utilized in this computer program) helped

create a data base of performance characteristics for several propellers, but it

did not give any indication as to how these theoretical results compared to

experimental results. A comparison was made using experimentally derived

performance data available for a 13-7, 14-4, and 14-8 propeller. The

geometrical parameters were measured for these propellers and input into

the performance program to obtain theoretical results. A study of the

experimental and theoretical results for the same propellers yielded a definite

trend. In general, the theoretical results for the power coefficient were much

lower than the "real" experimental" results. On the other hand, the

theoretical results for the thrust coefficient were much higher than the

experimental results. In order to correct the theoretical data, a quantification

PropulsionSystem 4 -3
of these differences was needed. Correction factors as a function of advance

ratio (J) were ascertained for both the power coefficient (Cp) and torque

coefficient (Cq). This correction function incorporated the average differences

between the theoretical and experimental results for both the thrust and

power coefficients over a range of advance ratios from 0.2 to 0.8. This

correction function was then added to the theoretical performance

characteristics to yield near "real" data. Figures 4-1 and 4-2 compare the

theoretical performance curves with the corrected performance curves for the

10-6 propeller. Table 4-2 gives the form of the correction function and the

appropriate coefficients.

TABLE 4-2: Correction Functions for Propeller Performance

Functional Form : CF = CO + CI*J 1 + C2*J 2 + O*J 3

Thrust, Ct : CO = -0.035761 C2 = -0.082838

C1 = 0.085253 C3 = -0.039541

Power, Cp : CO = 0.022765 C2 = 0.32328

C1-- -0.14785 C3= -0.2654

Effidency, q : CO- 0.83487 C2= 19.653

C1 _- -7.4942 C3 = -17.739

CF is added to theoretical value; J is advance ratio

In correcting the theoretical data, it was assumed that the correction factor was

the same for all of the propellers. Experimental results were not available for

smaller diameter propellers to validate this assumption. However, the

corrected performance data represented a significant improvement in regards

to accuracy over the theoretical data. In later analysis, the theoretical

performance curves were augmented by these correction functions. This

greatly enhanced the accuracy of the propeller selection process.

C. Battery / Propeller Selection

The next task in sizing the propulsion system was to determine the type

battery pack and propeller to be used with the Astro 05 motor. These two

Propulsion System 4 - 4
Ct

Figure 4-1

Thrust coefficient vs. Advance Ratio

0.20

0.16

0.12

0.08

0.04 "

0.00

-0.04

0.2

• I • I " I " I " I "

0.3 0.4 0.5 0.6 0.7

J

0.8

J Theory

Theory with corrections

Cp

Figure 4-2

Power coefficient vs. Advance Ratio

0.06

0.05"

0.04

0.03

0.02

0.01

0.00

-0.01

0.2

I I I I I

0.3 0.4 0.5 0.6 0.7

&

J

Theory

Theory with correcetions

(10-6)

0.8
propulsion system components were not independent of each other, rather

significantly related. Therefore, the process became one of considering

numerous battery pack/propeller combinations. Considered were battery

packs ranging from four to ten 1.2 volt Nickel-Cadmium batteries. The

battery type would be determined later from a consideration of milliamphours needed for endurance purposes.

In selecting the battery pack/propeller combination the three crucial phases of

flight -- namely, take-off, climb, and cruise - were examined. Each of these

flight phases were analyzed separately in order to select the battery/propeller

combination to yield the best overall performance.

Several constraints governed the selection of the propulsion system. These

are summarized in Table 4-3 and are discussed throughout this section in

regards to the different components and the processes by which they were

selected. Firstly, the maximum distance allowed for ground roll on take-off

was 75 feet and was set by the Design Requirements and Objectives study. The

maximum current allowed in the circuitry was to be 20 amperes. This arose

from the fact that the Astro 05 motor would be damaged if the armature

current exceeded this value and it was "enforced" by inserting a 20-amp fuse

into the circuitry so that it would blow and safe the motor. It was then

obviously undesirable to choose a battery pack/propeller combination that

would draw a current exceeding 20 amps at any time during flight because

this would cause an immediate loss of all power. Another constraint was a

minimum power available output of 37 Watts. This would ensure that the

plane climb at the desired rate of climb.

TABLE 4-3 : Constraints Governing Propulsion System

Maximum Take-Off Distance

Maximum Current Draw

Minimum Power Available

Maximum Battery Pack Weight

Maximum Endurance

Maximum Range

75 feet

20 amps

37 Watts

11 o_

78 seconds

1776 feet

Propulsion System 4 - 5
The selection of the battery pack and propeller was accomplished in two

related manners. One method examined the take-off phase and the associated

constraints. This method examined a wide range of battery pack/propeller

combinations and narrowed it down to one which was acceptable. The

second method examined the climb phase. It considered fewer propellers and

by applying relative constraints eliminated all but two combinations.

Considering results of both methods, the battery pack/propeller combination

was able to be selected with confidence.

Take-off Phase Method: This analysis proceeded with the take-off

performance of the airplane. This phase was aided by the computer program

"Take-Off Performance," written by Dr. S. M. Batill [Ref. 3]. It incorporated

the previously developed thrust and power coefficients as a function of

advance ratio, the lift and drag characteristics of the airplane, the wing area

and airplane weight, the ground friction coefficient of the mission flight

center, battery voltage and resistance, and the torque and current

characteristics for the Astro 05 motor. With these inputs, the program

integrated the ground roll performance at 0.05 second intervals and yielded

the following output: the distance required to achieve the flight speed;

maximum current draw; maximum power required to turn the propeller;

time to achieve take-off; and battery power consumed in ground roll. The

final selection was based upon take-off distance and maximum current draw,

even though we desired that the maximum battery drain not to exceed 5% of

the total milliamp-hour capacity. A sample of an input and output data file

are presented as Item 9 in the Appendix.

Using the propeller data, each propeller was analyzed for take-off

performance in conjunction with 5, 6, 7, and 8 batteries, with each battery

representing 1.2 volts. Although the amount of batteries used would have an

effect upon the battery pack weight, the weight of the airplane was assumed

to remain constant at 3 pounds. This weight estimation is based on 7

batteries, and hence if the number of batteries were less, resulting in a

decrease of total airplane weight, the overall take-off performance would

improve. This assumption was validated through use of the program and

initial hand calculations.

Propulsion System 4 - 6
Figures 4-3 through 4-5 represent the results determined through use of the

"Take-Off Performance" code. After analyzing the data, it was found that the

take-off distance and the maximum current draw defined a definite trend.

The grid upon which the curves are drawn represents all combinations of

propeller pitch and diameter studied. Each intersection point indicates one

such combination with the abscissa indicating the propeller diameter while

the ordinate is the propeller pitch. As indicated on the grid, several

combinations were eliminated because they did not overcome friction, takeoff in under 300 feet, achieve take-off velocity, or produce enough power.

These are indicated by the asterisks (*) according to the key at the bottom of

the figure.

The take-off distances formed parabolic curves that opened to the right

(Figure 4-3), while the maximum current draw formed parabolic curves that

opened to the left (Figure 4-4). This trend in data existed for all of the battery

pack sizes that were studied. In addition to these results, the battery drain in_

mahrs is represented for the different propeller sizes (Figure 4-5). The data

does not form any analytical trend because the battery drain is a function of

current draw as well as total time for take-off.

Referring to Figure 4-3, it is shown that for 7 batteries, there existed the

greatest number of propellers which met the take-off distance constraint.

These being the 10-6, 10-8, 11-6, 11-8, 12-4, and 12-6, all of which yielded a takeoff distance of 50 feet or less. Figure 4-4, however, shows that only three of

these propellers fall below the maximum allowable current draw, these being

the 10-6, 10-8 and 11-6 drawing approximately 15 amps, 19 amps, and 19 amps

respectively. The 10-8 and 11-6 were rejected because they were considered to

be within the limit of uncertainty associated with the accuracy of this analysis.

Hence, this left only the 10-6 propeller to be examined. The 10-6 yielded a

take-off distance of approximately 45 feet which fell well below the maximum

allowable take-off distance of 75 feet. In addition to having a relatively short

take-off distance, this battery pack/propeller combination drew a maximum

of only 15 amps, 25% less than the maximum allowable current draw. It

should also be noted that this combination, as shown by Figure 4-5, only

consumes approximately 15 mahrs during take-off, which is quite affordable.

Propulsion System 4 - 7
Figure 4-3

Take Off Distane

7 batteries / 8.4

(ft)

volts

10  ̧

J_

O

,i,-a

,m

IZ.

8

6

2

7 8 9 10 11 12 13 14

Diameter

* Friction Exceeds Stalic Thrust

*° Max Power Exceeds Max Allowable
Figure 4-4

Maximum Current Draw (Amps)

7 batteries / 8.4 volts

J_

Qg

a.

lo

8

4

\

\

\ \

"\ \

\

7 8 9

0 11

Diameter

2 13 14

• Friction Exceeds Static Thrust

•* Max Power Exceeds Max Allowable
Figure 4-5

Battery Drain (mahrs)

7 batteries / 8.4 volts

p.

o

lu

Q.

10"

8

6

4, ....

10 11 12 13 14

Diameter

" Friction Exceeds Static Thrust

** Max Power Exceeds Max Allowable
In summary, the take-off performance analysis indicated that the best

propeller/battery combination suited for take-off, allowing the airplane to

take-off in under 75 feet and drawing less than 20 amps, consisted of the 10-6

propeller.

Climb Phase Method : Initial calculations showed that the climb phase

proved to be the most demanding in terms of power requirements. From

estimations of cruise power requirements and desired rate of climb, we were

able to determine the amount of available power needed to be supplied to the

aircraft in order to carry out a successful climb. This needed power available

became one constraint which the battery pack/propeller had to fulfill.

Another constraint came from the allowable armature current. As

mentioned earlier, the Astro 05 motor would be damaged if it experienced a

current of greater than 20 amps. For the climb and cruise analysis, an

Airscrew 9-6, a Tornado 10-6 and a TopFlight 12-6 were considered so as to

represent the desired range of propellers from which to choose.

A plot of power available against the number of batteries for the three choices

of propellers is shown in Figure 4-6. From this figure, several combinations

are immediately ruled out since they do not provide the 37 watts of available

power needed in climb. The Airscrew 9-6 propeller with a battery pack of 7 or

fewer batteries fails to produce enough power. Likewise, the Tornado 10-6

with 6 or fewer and the TopFlight 12-6 with 5 or fewer batteries do not meet

the power constraint. These combinations were therefore eliminated as

possibilities. The armature current constraint served to exclude several other

possibilities. Figure 4-7, which shows the relationship between armature

current and the number of batteries for the propellers under consideration,

illustrates these excluded combinations. The Airscrew 9-6 propeller with a

battery pack of 10 batteries violates the current constraint. The combinations

involving the Tornado 10-6 with 10 and 9 batteries exceeded the allowable

current and were ruled out. This same propeller with 8 batteries was also

ruled out since the current draw of 19 amps was, as mentioned earlier, within

the limit of uncertainty associated with the accuracy of this analysis. The

combinations of the TopFlight 12-6 propeller with 10 to 6 batteries were also

eliminated from consideration due to the excessive current draw.

PropulsionSystem 4 -8
Figure 4-6

Power Available Variation with Batteries

120

100

e

80

> It

3= 60

0

Q.

20

0

2

• TopFlight 12-6

• • Tornado 10-6

[] • [] Airscrew 9-6

IN • []

[] • []

PA needed

.............. II"''''_'''11" ............................ in climb

[] • []

• I • I • I • ! , l

4 6 8 10 12

Number of Battelle•

e

._,-q

tO w

Q.

E

Q I

s_

E

Figure 4-7

Armature Current Variation with Batteries

4O

30

20

10

0

2

.. fuse limit

[] TopFlight 12-6

[]

[]

• Tornado 10-6

• •

....... 111......... v ......... .m Airscrew 9-6

II

e

u u

• •

• • el

a =

n , I i I , l n I

4 6 8 10 12

Number of B=,ttedes
Summarizing the results of these constraints, only two battery pack/propeller

combinations satisfied the requirements. These were the Airscrew 9-6 with 8

batteries and the Tornado 10-6 with 7 batteries. The former was eliminated

through a take-off analysis since it was unable to generate enough thrust to

exceed static friction. This enabled selection the Tornado 10-6 propeller and a

battery pack of 7 batteries to be used along with the Astro 05 motor.

The propulsion system had then been specified with the exception of two

parameters. These were the battery type (AAA, AA, C, or D) and the manner

in which the batteries were arranged (series or parallel). The factors

influencing the battery type were capacity (in terms of milliamp-hours) and

weight. This discussion will be presented later. The reasons for which the

batteries were arranged in series will now be discussed. Since the power

produced by the motor is directly proportional to the open circuit voltage it

was advantageous to arrange the batteries as to achieve a high voltage. This

was accomplished by a series configuration since voltage adds in series.

Secondly, the current does not add in series as it would in parallel. Therefore,

a series arrangement would keep the current low which is a desirable feature.

For these two reasons, the batteries were arranged in series.

Considering the Tornado 10-6 propeller with 7 batteries, Figure 4-7 shows that

the current draw during climb at full throttle is approximately 16 amps. For a

4 second climb, the power consumed is approximately 20 mahr, which is quite

affordable. Upon reaching a design altitude of 20 feet, the motor was throttled

back (by adjusting the voltage with a speed controller) so that the power

available would be equivalent to the power required. The power required

referred to is that which corresponds to a steady-level turn with a bank angle

of 30 ° . This is so that the airplane would not lose altitude in the turn. Figure

4-8 shows the relationship between armature current and power available for

a range of throttle settings. The power required for a steady-level turn was

about 8 watts. This figure shows that the corresponding current draw is

roughly 6 amps. Although the power required in steady, wings-level flight is

slightly less than that required for a steady-level turn, the two powers will be

considered equal so as to simplify the calculations and produce a conservative

estimate. Considering the power consumed in take-off (15 mahr), the power

Propulsion System 4 - 9
Figure 4-8

Relation Between Armature Current and Power Available

For Range of Throttle Settings

,.1u

0._ 2o

..........

........................... Bill ;"=_" -Ai

 oma o =

I

E 10 __Ji]... , ......

0

0 10 20 30 40 50 60 70 8()

Power Available

[ Watts ]

Figure 4-9

Relationship Between Throttle Setting and Armature Current

with Tornado 10-6 Propeller

1.:_ [- number of batteries

r 6 7 8

1.01- • • n

/ B • U

k m • n / i • n

= 0.8 r- m= • e,, u

t, r ==. ", °

0.8I- ==..= .

• I, men

o 0.4

_- 0.2

0.0

0 4 8 12 16 20

Armatui, e Current

[ amps ]
consumed in climb (20 mahr), the current draw in cruise (6 amps), and the

aforementioned simplification, the endurance of the aircraft is 221 seconds for

a battery pack of AA batteries and 129 seconds for 2/3AA batteries. The AA

batteries were chosen because of the large cushion in capacity. Although the

2/3AA batteries provided enough power capacity for the anticipated mission,

there were still unknowns that had to be considered. The larger capacity AA

batteries were chosen because it provided a margin which would allow for an

aborted take-off, ground roll testing prior to flight, and a safe margin to

account for standard unknowns. This comfortable increase in capacity with

the AA batteries also meant an increase in weight, from 2.94 to 5.46 oz.

Weight estimations for the initial design included a battery pack weight of 9-

11 ounces. All other parameters in the aircraft design were able to be

achieved assuming this battery pack weight. Therefore, by using the seven

AA batteries the weight was already well below initial estimations. Though

the additional weight savings with the 2/3AA batteries would have been

attractive, it was regarded as unnecessary in light of this already "low" battery

pack weight achieved with the AA batteries. The small weight penalty was

willing to be accepted in light of the increased endurance.

D. Speed Control

As mentioned previously, the propulsion system incorporated a speed

controller in order to throttle back the motor for the cruise portion of flight.

The take-off and climb phases of flight were performed at full throttle to

ensure the desired performance. Throttling back for cruise allowed the pilot

to maintain steady level flight through the turns as well as reducing battery

consumption in order to increase endurance. Due to "the small margin for

error in throttling back for cruise, it was desired to obtain some idea of the

degree of throttling necessary so that the pilot might be informed. Figure 4-9

shows the relationship between throttle settings and armature current for the

chosen Tornado 10-6 propeller. As revealed by the figure, for a cruise

armature current of 6 amps (with 7 batteries) the throttle setting was to be

approximately 50%. With this information, the pilot knew, upon reaching

cruise altitude, to throttle the speed controller to 50% power. From there,

Propulsion System 4 - 10
only minor adjustments were required to maintain a steady-level turn and

then steady-level flight.

Propulsion System 4 - II
SECTION V

STRUCTURES
A. Structural Configu.ration

Before estimation of the center of gravity could be completed, the structural

configuration of each component was determined. The structural design was

based on existing plans of other flight worthy aircraft found in modelling

magazines and other such literature. Regular design status reports by the

other design groups and aircraft design expert Mr. Joe Mergen also proved

quite helpful in determining the structural layout for the SCREEM_-J4D.

Fuselage : The fuselage is mainly a truss structure with minimal planking

to keep the weight low and to provide added strength. The structure is

shown in Figure 5-1. It is comprised of primarily balsa wood with birch

plywood used as mounting plates. The fuselage is 37" in length and has a

maximum cross section of 4"x 4". The main section of the fuselage was

designed in such a fashion as to allow for flexibility in placing the internal

components. This was done so that it would be easy to adjust the

components if a center of gravity shift was needed.

Wing : The wing is pictured in Figure 5-2. It consists of 3 sections: one 36"

inboard and two 30" outboard sections. The wing was designed in this way

simply for storage purposes (within a 2'x2'x4' compartment). The design is

fairly conventional. False ribs located at the leading edge are used for added

support for the skin. The analysis for the spar configuration is discussed in

the next section. The main features of the design are the connection of the

inboard and outboard wing, and the mounting location of the wing to the

fuselage. Initially, a 'micro-bubble/epoxy' matrix with an aluminum or brass

rod set in the spar box was considered for this purpose. But a simpler spruce

connection was decided upon for lighter weight and ease of construction. 4"

spruce beams will be set in the 'box' formed by the spars and webs (see Figure

5-3), one in the inboard section and one in the outboard section. The

connection will be secured together with tape wrapped around the wing.

Mounting of the wing onto the fuselage will be accomplished using the bolt

technique. Bolt locations are at the forward and trailing edge of the wing

center. Added support will be provided by securing the wing with a rubber

band, if needed.

Structur_ 5 - 1
b.. I,,. I,,I

.l_wI

•i._, Z

H

k

' l_, ̧

O

.b-,

Figure 5-1

Fuselage Structural Configuration

o_,,4

¢d3

/ I

__L-!

_-I l-;

\\ x..\

\\

x\\ x .. • "N

O

O
I

Figure 5-2

Wing Structural Configuration

•r .r ._

000

_00

_qr

_qrqr

* p,,,l

I

I

.

. _° w _ ---

0

,L

m

!

I

N

I

I

e:_

o

0
000

_ ¢,q q,

t4_4_4

1 fl

i !

Figure 5-3

Wing Connection

./

I

r

L_

¢

i
Empennage : The empennage (vertical and horizontal tails) are pictured in

Figure 5-4. The design is truss-like to provide strength and low weight. The

empennage will be permanently fastened to the fuselage.

B. Spar Configuration

Before analysis of the spar was performed, the spanwise distribution of shear

and bending moment for the cruise flight condition was determined. This

was done by means of the computer program listed as Item 10 in the

Appendix. Input for the computer code included:

a. Wing planform geometry

b. Aerodynamic section properties as a tabular function of spanwise

position

c. Structural weight per unit span as a tabular function of spanwise

position

d. Aircraft dynamic pressure and angle of attack

which are read from a data file. A second computer program listed as Item 11

in the Appendix was utilized to determine the centroid location, section

properties relative to the centroid, and the direct stress due to both bending

and axial loads for the idealized airfoil section shown in Figure 5-5. Each

component of the airfoil was idealized as a "lumped area". The maximum

direct stress, (3 = +-779 psi, was located in the main spars (areas 3 and 8) which

are spruce. The maximum allowable stress for spruce is 6200 psi which shows

that the main spars are safe from failure. Even when considering a factor of

safety of 2.0, the maximum direct stress is _ = 1560 psi, which is still safe from

failure. Table 5-1 summarizes the section properties, loading conditions,

direct and allowable stress for each "lumped area".

Structur_ 5 - 2
LaJ _a _

-'r .T. _

000

*.'* 00

_4_c4

Figure 5-4

Empennage Structural Configuration

_i t .>!-

Vertical Tail

Horizontal Tail

t I
i_ tot ult

000

*nO0

*.* f_ qr

qr qr qr

w. jB um

Figure 5-5

Spar Configuration

I --_Ic_ _ Lm O -.T _ cN um

I I I I

• ,"_ C 0 u"_ 0 0

I I I I

• QQOOO000
Figure 5-6

/

[

f

/

!

I

/

L

4

\

\

Spar Configuration

/I e*leoe

II

_gooooo 00000

_ I.eiJo

II I

__o

oooo_

e.ee.e

< oooooo
TABLE 5-1 : Direct Stress Distribution at Wing Root

Ycentroid ffi 0.(30 in

Zcentroid ffi-2.81 in

Iyy = 36.59 in 4

Izz ffi 0.42 in 4

Iy z = 034 in 4

P ffiO.O01b

My ffi 2.42 in-lb

Mz = 30.24 in-lb

i _xx(i) psi ¢_xx(all) psi Material

1 -17.28 +400.00 Balsa

2 0.09 _+400.00 Balsa

3 -779.12 _+6200.00 Spruce

4 -0.09 _+400.00 Balsa

5 -36.62 _+400.00 Balsa

6 -1.96 +400.00 Balsa

7 18.37 _+400.00 Ba Isa

8 779.12 +__6200.00 Spruce

An analysis of the joints connecting the inboard and outboard wings was also

performed. Figure 5-6 shows the configuration of the spars and Table 5-2

summarizes the data. As the table indicates, the maximum allowable stress is

not reached in any of the members.

TABLE 5-2 :

Ycentroid = 0.12 in

Zcentroid = -0.99 in

Direct Stress Distribution at Y=18 in

Iyy = 13.14 in 4

Izz = 0.44 in 4

Iyz = 0.11 in 4

P = 0.00 Ib

My = 0.85 in-lb

Mz = 10.60 in-lb

1 -258.07 +6200.00 Spruce

2 -12.20 +400.00 Balsa

3 -7.64 +400.00 Balsa

4 258.07 +__6200.00 Spruce

5 -91.08 +__6200.00 Balsa

6 -61.78 +6200.00 Balsa

Structures 5 - 3
C. Materials Selection

The critical characteristics of the materials under consideration were:

a. Strength

b. Weight

c. Cost

d. Machinability

e. Availability

Wood, metal, composites, and ceramics were considered and the relative

rankings of these regarding the critical characteristics are summarized in

Table 5-3.

TABLE 5-3 : Relative Material Characteristics

Material Stren k_h Cost Weight Machinability Availal_ility

Wood 4 1 2 1 1

Metal 2 2 4 2 2

Composites 1 4 1 3 3

Ceramics 3 3 3 4 4

The table shows that wood is the best material to be used. Though wood

ranked last in strength, it still would provided sufficient strength to meet the

stress requirements mentioned earlier. The bulk of the structure was

comprised of balsa wood with spruce and birch plywood used where extra

strength was needed. The skin of the J4D was "Micafilm" manufactured by

Coverlite. It is a non-woven fiber that "offers low weight, resistance to

tearing, and good modulus" fit for models where weight control is a primary

concern.

D. V-n Diagram

The V-n diagram is shown in Figure 5-7. First notice that the gust loads do

not affect this diagram because of the indoor operating environment which is

free from any significant gusts. There are no negative load factors because the

aircraft is not designed to operate at a negative wing incidence. The

maximum operating velocity, Vmax, of the aircraft is 58 ft/s, the minimum

Structures 5 - 4
Flgure 5-7: V-n Dlagram

.

ult', ................................................................

Iw

0

0

Ca

tL C

"0 "

¢0

0

.J

=_E .-_=

_>

-I

Veloclty (ft/s)
(stall)velocity, Vrnin is 18.9 ftls,and the cruise velocity, Vcruise,is23 ft/s. The

limit load factor, nlim, is 1.15 (from a maximum bank angle of 30 degrees).

The positive stalling portion of the curve is found using standard sea-level

conditions (p = 0.00237 Ib/fta),a maximum liftcoefficient of 1.3 and a wing

loading of 0.55 Ib/ft2. The ultimate load factor,nult = 1.73,is based on a factor

of safety of 1.5 (nult= 1.5nlim). Special attention should be made to the fact

that the cruise point (n=l.0, V=Vcruise), pushes the limits of the flight

envelope. Therefore, the aircraftmust not be flown at an attitude such that

disastrous loads are placed on the aircraft. The factor of safety,however, does

allow for somewhat significantloads to be placed on the aircraft,ifnecessary.

TABLE 5-4 : Special Flight Conditions Illustrated in V-n Diagram

Vmax 58 ft/s

Vmin 19 ft/s

Vcruise 23 ft/s

CLmax 1.3

nlim 1.15

nult 1.73

_rnax (bank angle) 30 degrees

W/S (wing loading) .00237 Ib/ft3

.0379 oz/ft 3

F.S. (factor of safety) 1.5

Structures 5-5
SECTION VI

WEIGHTS ANALYSIS
A. Weights Estimation

An estimation of component weights was initially made based on standard

component weight fractions for remote control aircraft of a similar type.

These provided the initial goal weights for the components which guided the

design of the aircraft. Based on an overall "target weight" for the aircraft of 3

pounds (48 ounces), the component weight fractions were determined. These

are summarized in Table 6-1.

TABLE 6-1: Initial Estimation of Component Weights

Com_nent Weight (oz) Wei[_ht Fraction (%)

Structural Frame 19.2 40.0

Propulsion 17.3 36.0

Avionics 7.2 15.0

Landing Gear 2.1 5.0

Pa_,load 2.0 4.0

After the propulsion system, wing, control surfaces, and fuselage were sized

and the material selected, a more refined estimation of component weights

was made. The structural weight (fuselage, empennage, wing) were estimated

from the structural configuration discussed in Section V. The propulsion

system analysis resulted in the selection of the motor, batteries and propeller,

and hence the propulsion weight. Avionics weight was known from the

knowledge of the servos and speed controller used. The payload was

designated to be 2 ounces. The landing gear remained a true estimation since

its design had not been determined. The refined component weights and

weight fractions are summarized in Table 6-2.

Weights Analysis 6 - l
TABLE 6-2: Refined Estimation of Component Weights

Co_t Weight (oz) Weight Fraction (%)

Structures 21.26 44.3

Wing 10.36 21.6

Fuselage/Empennage 6.4 13.3

Skin, Fasteners, Links, etc. 4.50 9.4

Propulsion 16.19 33.7

Propeller 1.00 2.0

Fuel 5.46 11.4

Engine 9.73 20.3

Avionics 4.15 8.6

Landing Gear 2.40 5.0

Payload 2.00 4.0

The avionics components consisted of the receiver, system batteries, and

servos, while the engine consists of the motor and speed controller. The

"skin, fasteners, links, etc." component is that material which remained for

use in construction of the prototype.

B. Internai Layout

In determining the internal arrangement of the aircraft, several factors had to

be considered. Those factors of influence were:

- effects on center of gravity location

- proximity requirements

- simplicity of layout: non-interfering components

- cooling paths

- access

- wing mounting space designation

With the above considerations, the internal layout was determined.

However, the locations were able to be changed slightly in the event that

center of gravity location of the prototype was different from the designed

Weights Analysis 6 - 2
location. That is, there was some flexibility in this layout to accommodate for

the uncertainties of the construction phase.

Table 6-3 is a summary of the internal system components, their respective

functional classifications, and the corresponding sizes and weights:

TABLE 6-3: Weights and Sizes of Internal Components

Com_xment Classification W [oz.] L [in] w [in] h [in]

Motor propulsion 6.5 2.25 1.25 **

Battery Pack (7)" power supply 5.46 1.95 0.55 **

Sys. Battery(4)* power supply 2.0 1.73 0.39 **

Receiver control system 0.95 1.31 1.87 0.81

Servos (2)* control system 0.6 1.06 1.5 1.12

Speed control control system 3.23 1.875 3.0 1.25

Payload payload 2.0 2.0 2.0 2.0

value in parenthesis indicates number of such component

* weight and sizes data is given per battery or servo

*" indicates diameter

Note that the battery pack is comprised of seven AA batteries, each of which is

1.95 inches in length and 0.5 inches in diameter. Their combined weight, as

indicated above, is 5.46 ounces. Also, the system batteries were AAA, 1.73

inches long and .39 inches in diameter. Note that there are two servos: one

operates the rudder surface and the other operates the horizontal tail surface.

To determine the relative location of the internal system components, a

weight distribution program was utilized (see the computer program listed as

Item 12 in the Appendix). With fixed values for the structural center of

gravity, the placement of the internal components was selected. With the

location of the components already set, the program was used to determine

the location of the battery pack (whose weight was considered to be ballast).

As mentioned before, the location of the battery pack and other components

were allowed to shift slightly so that the overall center of gravity would

remain in the design location.

Weights Analysis 6 - 3
The design locations of the internal components are summarized in Table 6-4

and shown in Figure 6-1.

TABLE 6-4: Design Locations for Internal Components

Component Center of Gravity Location [in]

x z _,

motor 1.13 2.63 2.0

speed controller 4.0 2.63 2.0

receiver 6.0 0.63 2.0

system battery 6.0 2.28 2.0

servo 1 (rudder) 18.94 2.44 1.0

servo 2 (horiz. tail) 18.94 2.44 3.0

battery pack 9.5 1.5 2.0

payload 9.0 3.0 2.0

x is measured from the nose of the aircraft, increasing toward the tail

z is measured from the bottom of the fuselage, increasing toward the top

y is measured from the middle of the fuselage, increasing toward the right

The servos are placed in the rear so there is no interference with the control

lines to the rudder and elevator. They will be mounted on 1/16" balsa wood

strips (servo trays), using wood screws. All other internal components, with

the exception of the motor and the battery pack, will be similarly mounted.

To avoid wood splitting at the screw holes, epoxy will be applied in the holes

prior to inserting the screws. The motor will rest on two triangular, balsa

wood mounts, and will be secured by elastic straps and screws. The battery

pack will be secured by Velcro straps, to allow for quick re-location in the

event that additional modifications are made which affect the center of

gravity location. Thus, the battery pack provides a means for center of

gravity travel.

Weights Analysis 6 - 4

C. Center of Gravity Estimation

The placement and location of the center of gravity of each component are

shown in Figure 6-2. The CG of the overall structure is set at 33% of the chord

of the wing and the locations of the vertical and horizontal tail leading edges

are set at 31.7" and 33.2", respectively. With these values fixed, the placement

of the internal components was selected as outlined in the previous section.

Weights Analysis 6 -5
Lia _aa

_m

00

O0

!

Figure 6-2

Center of Gravity Estimation

!

I

i

I

A

T

t _

i

CO

,r

i

i

I

t
SECTION VII

PERFORMANCE
A. Lift, Drag and L_/Drag

As for every aircraft, the basic size, weight, and cruise velocity of the aircraft

dictate the lift coefficient necessary to sustain steady, level flight. The

SCREEM-J4D, with its planform area of 5.46 ft 2, weight of 3 lbs, and cruise

velocity of 23 feet per second needs a CL of 0.874 tO maintain steady, level

flight. The lift coefficient necessary for flight is plotted as a function of

velocity on Figure 7-1. The drag breakdown described in Section II gives a

drag polar of CD = .0319 + CL 2/.682/I;AR. Using the SCREEM-J4D's aspect ratio

of 11.72 and cruise CL of .874, the drag coefficient at cruise is then .0564 (see

Figure 7-2) for a total drag of .194 pounds. Combining this drag and the cruise

velocity, the power required at cruise is 6.05 Watts. While this is not the

minimum power required for this aircraft, (which incidentally occurs at 16-18

feet per second, as can be seen in Figure 7-3) it is only 25% higher than this

minimum power required. With the lift and drag coefficients thus calculated,

the maximum lift/drag ratio is 14.0, and as is evident in Figure 7-4, the

SCREEM-J4D is flying at maximum lift to drag ratio at cruise.

B. Take-Off and Turning Hight

The SCREEM-J4D produces a static thrust of 1.186 pounds, which allows for

takeoff in 45 feet at a takeoff velocity of 23.8 feet per second and CL )max of 1.2.

The power required to climb at the prescribed rate of 6 feet per second is 37

Watts, and the relatively high power available of 50 Watts means that the

SCREEM-J4D will possess the capability to climb at a much higher rate of 15.2

feet per second.

In turning flight at the prescribed turn radius of 40 feet, the SCREEM-J4D

must bank to an angle of 23 °, which means that the load factor in the turns

rises to 1.1. At a roll rate of 28.8 degrees per second, banking to this angle

requires .8 seconds and 18.4 feet to complete the transition to "fully banked."

In order to perform this maneuver, the SCREEM-J4D must induce a change

in angle of attack on the inboard wing of -3 ° and a change of +3 ° on the

Performance 7 - 1
W

.J

5.0

2.5

Figure 7-i

8.8

18 IS a8

UELOC I TY £+"t/s ]

aS 38

1.e

e.S

e.O

10

Figure 7-2

IS as

UELOC I TY [ f"t/s ]

2S 38
O.

1a.5

10.0

10

Figure 7-3

15 a8

UELOCITY El"t/s]

0

1-4

p-

<Z

n,

i5

,%

0

ppU.

1--4

.J

15

10

5

Figure 7-4

10 15 aS

VELOCITY [¢t/s]
outboard wing. This change is incurred by a maximum rudder deflection of

25 ° . Reducing the rudder deflection will reduce the roll rate, but even cutting

the roll rate in half will not severely inhibit the roll performance of the

aircraft.

C. Endurance and Range

Given the battery capacity of 500 mahr the SCREEM-J4D, the maximum

endurance is 221 seconds. Broken down into the different sections of the

flight, 15 mahr is used in the takeoff phase and 20 mahr is used during the

climb phase, leaving 465 mahr for the steady, level phase of the flight. This

endurance should be sufficient for the mission, which will only require 78

seconds. Combining this endurance with the cruise velocity of 23 feet per

second gives a maximum range of 5500 feet, once again well above the

required range of 1800 feet. Parts B and C of this section are summarized in

Figure 7-5 while Table 7-1 summarizes all of the critical performance data.

TABLE 7-1 : Critical Performance Data

Vcruise 23 ft/s n turn

CL in cruise .874 p roll rate

CD in cruise .0564

L/D max 15.5 E max

Pr in cruise 6.05 Watts R max

Pa 50 Watts "_min glide angle

Pa needed for climb 37 Watts R/D

R/C max 15 ft/s

1.1

29 °/sec

for6r= 25°

221 sec

5500 feet

4 degrees

1.6 ft/s

Performance 7 - 2
Figure 7-5

III

M

),

Ib

@

)

@

a.

Power Profile for Final Design

O0

design

I _ velooitg ....

60 -I _ _ ; Power Avai,ao,e

i

; Power Required

20 _

0

0 I0 20 30 40 50

Velocitg [ ft/s ]

Motor:

Propeller:

gaLLeries:

Astro 05

Tornado 10-6

7 AA Nickel-Cadmium

Endurance:

Range:

R/C:

T-O Distance:

221 seconds

5500 FeeL

6 ftls

45 FeeL
D. t ndin8

Because the aircraft will be flying at its maximum lift to drag ratio at cruise, its

glide angle will be maximum at cruise. This maximum glide angle is -4 ° ,

which will give a minimum rate of descent of 1.6 feet per second. At this rate

of descent, the descent phase will take approximately 12.5 seconds, requiring

close to 300 feet from the start of the descent phase until touchdown. This

distance will be shortened somewhat by descending at a larger angle, with a

flare upon touchdown to reduce the impact force. Due to the rough surface of

the runway, it is estimated that for a touchdown on the first half of the

available runway, the aircraft will come to a complete stop by the end. As a

precaution, a net will be available after landing if it is evident from flight tests

that the plane requires too much distance to stop.

Performance 7 - 3
SECTION VIII

TECHNOLOGY

DEMONSTRATOR
A. Construction Phase

The product which emerges from the manufacturing phase is rarely identical

to the product of the design phase. The various problems and solutions

which were encountered in the construction of the SCREEM-J4D, as well as

the resulting variations from the design, are presented in this part.

The aircraft construction was facilitated by the detailed design studies,

especially regarding the structural framework. The fabrication process was

aided by actual size blue prints which served as templates for the assembly of

the fuselage, wing, and empennage. The structure was made of balsa wood,

spruce and plywood as designed and held together by epoxy and model glue.

The skin was a MonoKote plastic.

The weight of the prototype compared very well with the design weight. Each

component of the prototype weighed approximately as it had been designed

with the overall weight being .05 pounds less than was designed. This was

due to weight saving techniques employed during construction. Holes were

punched in non-critical structural components -- airfoils, false ribs, trailing

edge - and the empennage was built using fewer braces than designed to save

weight. The MonoKote plastic was also used on the fuselage to eliminate the

need for heavier balsa planking.

The prototype did differ from the design in several ways. First, the center of

gravity position was designed to be at 33% of the chord. However, the

prototype had its center of gravity at 40% of the chord. This was due in part to

the internal arrangement being slightly different. After the first flight test it

was decided that the CG had to be moved up to the 25% position. This center

of gravity shift was accomplished by lengthening the nose of the fuselage by

2.6 inches. This represented a significant difference from the design.

The chord of the main wing was to be 8.2 inches. However, the blue prints

were drawn with an 8 inch chord and the trailing edge piece was shortened to

adjust to the change.

Technology Demonstrator 8 - 1
The prototype possessed a greater endurance than the design. The design

batteries were not available for purchase. The batteries that were bought

differed only in capacity (700 mahr instead of 500 mahr); the voltage and

weight specifications remained the same. This change only affected the

endurance, which was significantly increased.

Another striking difference was related to the wing. The design called for the

wing to be in three sections with the outboard sections removable for storage.

The joint which connected the outboards to the main was not considered to

be structurally sound by an expert model builder late in the construction

phase. It was suggested that the joint be rebuilt for greater strength. It was at

this point that the executive decision was made to relax the requirement that

the aircraft fit in a 2'x2'x4' travel box, which restricted any piece from being

greater than four feet in length. This enabled AETC to permanently affix the

outboard sections to the main wing, greatly increasing the wing's strength

and producing a piece 8 feet in length.

B. Flight Tests

The SCREEM-J4D was constantly examined and tested during the

construction phase. The motor and control surface actuators were checked as

was the strength of the aircraft's structure. It was decided that before the J4D

flew its first test flight it should undergo an extensive ground check.

Described here are the pre-flight test plan and the actual test flight plan.

Pre-Flight Test Plan : The pre-flight test was intended to check all systems so

as to insure a successful flight around the required course. Those capabilities

which were examined were: control surface response; ground roll

controllability; power production; structural integrity; and lift generation.

Taxi testing was first performed to see how well the aircraft rolled down the

runway and was controllable by means of the rear wheel and control surfaces.

The power production was partially examined by measuring the speed of the

ground roll. The aircraft was held by the wingtips to test their strength.

Lastly, the aircraft was allowed to accelerate to take-off velocity and then

Technology Demonstrator 8 - 2
flown for a few feet very close to the ground to better understand its in-flight

behavior.

This test was carefully observed in order to identify problems which should

be addressed before the actual flight.

Flight Test Plan : The course to be flown is shown in Figures 8-1 and 8-2.

The course consists of three figure-8 maneuvers preceded by take-off and

climb and followed by landing within the shaded area. The ground roll

should cover no more than 75 feet with the aircraft becoming airborne by

point 1. The aircraft then climbs at a 15 ° angle until it reaches cruise altitude

of 20 feet at the point marked 2. It then banks left and begins its first figure-8

maneuver at point 3. Point 3a indicates the other end of the maneuver

whose radius of curvature at the ends is 40 feet. After three figure-8's, the

aircraft begins to fly a straight path at point 4, banks around the end of the

flight zone and begins decent at point 4a, touches down after a flare

maneuver at point 5, and comes to a complete stop by point 6.

The zone allowed for this flight is 300 feet long, 120 feet wide, and 25 feet

high. These dimensions are shown in Figure 8-1 and 8-2. The facility used

encloses a football field and has physical dimensions larger than the allowable

flight zone. This flight is flown in a controlled atmosphere with limited air

disturbances.

C. Instructions For Pilot

The pilot was informed of the capabilities and limitations of the aircraft in

regards to power, turning capability, bank angles, maximum angle of attack,

and cruise velocity. He was also told of the flight plan as described above.

The pilot will position the throttle such that the aircraft will roll at a modest

speed, but not take off. He will get a feel for the control surfaces as well as for

the turning capabilities of the rear wheel. After he is satisfied with the feel,

he will begin the mission flight.

Technology Demonstrator 8 - 3
Top View

Fisum _-1

i n i i

'-" 150 feet "-'

75 feet "-J

120 5_et

300 feet _='-I

Scale: I l = 30_et

View

_---" 4L--- _a.

_0-4- ,#_=--=--,---==-======o==.===.==..=.=,=.========.== ........ .=e===

I °*. ...... -" 2. " ..... @.

w-,- -"'-... s. I ..... - ..... " 4.

i ,I I _i I I | go " I I |, '-
The pilot knows that the maximum take-off distance is 75 feet, which is

marked on the runway. He will stand anywhere he feels comfortable and

begin the ground roll at full throttle. If for any reason the aircraft fails to

achieve flight by the 75-foot mark or it turns out of control, he will abort takeoff immediately. For the case of a successful take-off, he will climb at full

throttle to an altitude of 20 feet at which point he will throttle down to

roughly 50%. When the aircraft is over the end of the runway (even with the

pylon, at point 2) he will begin a left turn with approximately a 25 ° bank

angle. However, if for some reason the aircraft does not reach a 20-foot

altitude by the pylon, he should begin a turn anyway to avoid leaving the

allowable flight zone. In this case, he should maintain a full throttle setting

until he is able to reach 20 feet. If an altitude of 20 feet is unattainable, the

pilot should land the plane as soon as possible.

Upon successfully climbing and turning, the pilot will proceed to fly three

figure-8 maneuvers. The throttle setting for steady, wings-level straighf

portions of flight is approximately 50%. The pilot will increase the power

slightly in the turns so that the aircraft will not loose altitude in the turns. He

will fly the straight portions of the flight as accurately as possible so that

measurements may be taken by the design team.

If at any time the pilot feels uncomfortable with the qualities of the aircraft --

i.e. stall likely, dynamically unstable, difficult to maintain altitude,

unresponsive, ineffective controls, etc. -- he is advised to land it as soon as

possible in any safe location. The condition of the aircraft is of higher priority

than fulfilling the mission requirements. If the aircraft remains in good

condition, it may be modified to fly later and successfully meet the

requirements.

After completing the three figure-8's, the pilot will fly the aircraft around the

end, throttle down the motor, descend, and land at the beginning of the

runway. Upon touchdown, he will completely throttle back the motor and

guide it to a safe stop.

Technology Demonstrator 8 - 4
D. Safety Considerations

The SCREEM-J4D has not passed any official safety test, therefore it should

not be flown except by an experienced RPV pilot. All spectators should be a

safe distance from the flight zone and fully attentive in case of a "runaway"

aircraft. Members of the design team within the flight zone should be aware

at all times. No one should be at the end of the runway as the aircraft takes

off or lands or under the aircraft as it makes a turn. The aircraft may stall in

these circumstances and fall upon someone.

No one should be within three feet of the aircraft while the motor is running.

The propellers are extremely dangerous and often hard to see. No one should

ever touch, pick up, or work on the aircraft unless the battery pack is

disconnected from the motor.

During the design flight, NO ONE should be in any position which the pilot

has designated as unsafe.

The pilot must realize his restrictions flying indoors; hence, he must take

protective measures including possibly aborting the mission so as not to

jeopardize the facility and/or spectators.

Components inside the aircraft generate heat and may be very hot after

flight or taxi test. The aircraft should be handled with the idea that

some components may be of high temperatures. The aircraft should be

allowed to cool between missions.

Technology Demonstrator 8 - 5
F. Cost Analysis

The cost breakdown for the SCREEM-J4D is as follows:

Item or System Approximate Cost ($) Notes

Propulsion System

Motor & Gearing 75.00

Propeller 2.00

Batteries 15.75

Speed Controller 100.00

Avionics

Transmitter 200.00

Receiver 40.00

System Batteries 9.00

Servos 88.00

Actuation Rods 5.00

Hinges and Attachments 20.00

Structures

Balsa 40.00

Spruce 5.00

Glue 15.00

MonoKote skin 20.00

Labor 536.00

Administrative Costs 50.00

Total $1220.75

7 AA NiCad

4 AAAA NiCad

2

2 rolls

160 hours @ $3.75

cop_,in_ & binding

Technology Demonstrator 8 - 6
G. Critical Data Summary

The final form of the Critical Data Summary is

DESIGNGOALS:

ReCruise 100,000

V cruise 23 ft/s

Altitude cruise 20 ft

Turn radius 40 ft

Endurance 78 s

Weight 48 oz

BASIC CONFIG:

Wing Area 5.46 ft2

Weight 48 oz

Wing loading 8.79 oz/ft a

length 37.2 in

span 96 in (8 ft)

height 14 in

WING: 1 (low mount)

Span 8 ft

Aspect Ratio 11.72

Root Chord 8.2 in

taper Ratio 1

Airfoil section NACA 4415

t/c 15%

Incidence angle 4.5 degrees

C mac - MAC 8.2

Hor. pos of 1/4 MAC 9.017 in

Ver. pos of 1/4 MAC -4 in

Dihedral 10 degrees

Twist 0

Sweep 0

FUSELAGE:

Length 37. in

Diameter - max 4 x 4 in

Fineness ratio 9.3

Payload volume 25.17 cu in

included below.

EMPENNAGE:

Horizontal tail:

Area

aspect ratio

root chord

tip chord

taper ratio

I.e. sweep

incidence angle

hor. pos. of 1/4 MAC

vet. pos. of 1/4 MAC

Airfoil section

Vertical Tail:

Area

aspect ratio

root chord

tip chord

taper ratio

I.e. sweep

hor. pos. of 1/4 MAC

Airfoil section

AERODYNAMICS:

Cl max

lift curve slope

CDo

efficiency ( e )

Alpha stall

Alpha zero lift

L/D max

.625 ft 2

2.722 in

5.75 in

5.75 in

1

0

0 degrees

2.9 ft

0

flat plate

.38 ft2

1.38

7.2 in

3.9 in

0.54

23.2 degrees

2.749 ft

flat plate

1.2

3.72/rad

0.0319

0.682

14.5 (wing)

-4 (wing)

15.5

Technology Demonstrator 8 - 7
WEIGHTS:

Weight total

C.G. (x-pos)

Avionics

Payload

Engine

Propeller

Fuel

Structure

Wing

Fuselage/emp.

Landing gear

PROPULSION:

Type

number

placement

Pavil max @engine

Propeller diameter

Propeller pitch

Number of blades

STABILITY AND

CONTROL:

Neutral point

Static margin %MAC

Hor tail volume ratio

Vert tail volume ratio

Elevator area

Elevator max

deflection

Rudder Area

Rudder max deflection

Aileron Area

Aileron max deflection

Cm alpha

Cn beta

CI alpha tail

Cn delta r tail

Cm delta e

48 oz

.33c

7.38

2

6.5

1

5.46

10.36

5.22/1.18

2.4

Astro 05

1

front

100W / 50 W

10 in.

6 in.

2

.514 c

18.40%

0.356

0.016

.208 ft2

15 degrees

.21 ft

25 degrees

none

none

-0.685 [1/rad]

7.68 E-4

3.73

5.33 E-4 (neg)

-0.683 [1/rad]

PERFORMANCE:

Vmin

Vmax

Vstall

Range max

Endurance Max

ROC max

Min Glide angle

T/O distance

T/O rotation angle

Landing Distance

Bank Angle

Turn Radius

SYSTEMS:

Landing gear type

Main gear position

Main gear length

Main gear tire size

tail gear position

tail gear length

tail tire size

engine speed control

Control surfaces

TECH DEMO:

Weight total Take off

Weight empty

Wing Area

Hor. Tail Area

Vert Tail Area

C.G. posRion

1/4 MAC position

Static margin %MAC

V takeoff

Range

Endurance

V cruise

Re cruise

Turn radius

18.9 ft/s

58.6 ft/s

18.9 ft/s

5500 ft

3.9 min

6.3 ft/s

4.1 degrees

42 ft

11 degrees

50 ft

30 degrees

28.6 ft

tail dragger

wings

5 in

1 in diameter

rudder

.5 in

.5 in diameter

yes

Rudder/elevator

49 oz

49 oz

5.23 ft 2

.625 ft 2

.38 ft 2

.27 c

9.0 in

18 %

30 ft/s

8000 ft

240 s

33 ft/s

128,000

30 ft

Technology Demonstrator 8 - 8
SECTION IX

CONCLUSIONS
A. Disadvantages and Advantages

While the SCREEM-J4D will accomplish the proposed mission with a margin

of safety, there are a few short comings to the design. The allowable payload

is limited to a volume of 25.17 cubic inches and a weight of 2 ounces, to meet

our design requirements. The J4D has also been designed to meet the specific

requirements of this proposal. This greatly restricts the design flight envelop

as compared to general aviation RPVs. In addition to the restriction on the

flight envelop, the J4D was design to fly indoors, in a controlled

environment. While this will not restrict the SCREEM-J4D to indoor flights,

it has not been design to handle hostile environments, including wind gusts

and adverse weather conditions. This does not preclude the SCREEM-J4D

from all inclement weather; however, further analysis of its durability is

suggested.

The SCREEM-J4D is not without its advantages. Due primarily to the

constraints imposed on the design by AETC and the initial request for

proposal, the J4D has been designed as low speed, lightweight, highly

maneuverable RPV. In addition, it has been design for disassemble and

transportation within an 2 x 2 x 4 cubic foot volume. A further advantage of

the disassembly, is the ability to interchange parts. This facilitates the

replacement of broken or damaged parts. The disassembly will also aid in the

modification of the J4D for other missions.

B. Applications

From the analysis presented in this proposal, the SCREEM-J4D will perform

the the required station keeping mission within the design constraints as set

by AETC, with a reliable margin of safety. Due to the stringent constraints, as

set by AETC, the SCREEM-J4D can be a formidable option for numerous other

missions. Due to the goals of this mission, the battery pack and thus the

endurance are held to a minimum. Through an increase in the milliamphours, corresponding to an increase in battery size, the endurance of the J4D

can be increased to meet the requirements of a prolonged station keeping

mission. Unfortunately, the increase in battery size will increase the weight

Conclusio_ 9 - 1
and required cruise velocity. Prolonged station keeping mission may include

low altitude surveillance or electronic cloaking for military operations. Since

the SCREEM-J4D has been designed to perform its mission at low speeds, it

may be flown at speeds comparable to Navy ships. An electronic device may

then be loaded on board to emit artificial waves that simulate electromagnetic

emission of a Naval vessel.

Since the SCREEM-J4D is designed with a non-airbreathing engine, it may be

easily adapted to work as an atmospheric sampler, in both toxic and non-toxic

environments. Due to the design cruise velocity, the J4D may be modified for

function as a high altitude long duration station keeping mission. The

SCREEM-J4D was initially designed to complete a station keeping mission;

thus, it has been designed to optimize steady-state circling. The required

modifications for a prolonged flight mission were discussed above. Recall,

that these modifications will increase the cruise velocity. The cruise velocity

will be further increased for applications at high altitudes, due to a decrease in

density. This decrease in density will correspond to a decrease in lift for a

constant velocity. Thus, a high altitude mission will require an increase in

cruise velocity. While it may be necessary to select another engine to meet

these requirements, the present design only mandates 50% of the design

engine torque. This short discussion has presented a few of the numerous

missions that the SCREEM-J4D may be modified to accomplish

C. Final Word From The Design Team

This project has progressed rapidly from beginning to end, no doubt to the

dedication and incessant efforts of the team members. The design has

evolved from a rough sketch and a lot of hand waving into a polished aircraft

more than capable of performing the mission for which it was intended. The

SCREEM-J4D has a singular mission at this time; but its basic design, as we

have already seen, can be easily modified into one capable of performing a

number of far-reaching surveillance and reconnaisance missions. Its basic

airframe is lightweight and durable, lending itself well to effortless

transportation and on-site assembly. In addition to being extremely quiet, its

electric motor is perfect for missions in atmospherically hostile

Conclusions 9 - 2
environments. We, the members of AETC, support the SCREEM-J4D fully

and have no doubt that it will be successful in its intended mission and any

other applications that follow.

Eric T. Fick

Edward M. Dieser

Mt/chael J._ark

Chris "Duke" Sturgis

Robert M. CaUoway

Je/f_ey L_.janicik /

Stephen T. Valentine

Conclusions 9 - 3
REFERENCES
REFERENCES

1. ANDERSON, John D. Jr., Introduction to Flight, New York: McGrawHill Book Company, 1985.

2. ANDERSON, John D. Jr., Fundamentals of Aerodynamics, New York:

McGraw-Hill Book Company, 1984.

3. BATILL, Dr. Stephen M., "Take-Off Performance," Computer Code,

University of Notre Dame, 1990.

4. BERON-RAWDON, Blaine, Model Aviation, "Dihedral," four-part

series, August, September, October, and December 1989.

5. DUNN, Dr. Patrick F., "Design Rules of Thumb," University of Notre

Dame, 1989.

6. JENSEN, Daniel T., "A Drag Prediction Methodology for Low

Reynolds Number Flight Vehicles," University of Notre Dame,

January, 1990.

7. NELSON, Dr. Robert C., Atmospheric Flight Mechanics, University

of Notre Dame, Spring, 1989.

o

.

10.

11.

12.

NELSON, Dr. Robert C., Flight Stability and Automatic Control, New

York: McGraw-HiU Book Company, 1989.

STINTON, Darrol, The Design of the Aeroplane, New York: Van

Nostrand Reinhold Company, 1983.

USAF Stability and Control DATCOM, Flight Control Division, Air

Force Flight Dynamics Laboratory, Wright-Patterson AFB, OH.

YOUNG, Barry N., "Propeller Performance Analysis for Small

Computers," University of Notre Dame, May, 1984.

NATIONAL TECHNOLOGY INFORMATION SOURCE, "A Catalog of

Low Reynolds Number Airfoil Data for Wind Turbine Applications,"

College Station, TX, February 1962.
APPENDIX
3ALFORD UNIVERSITY FTN77 VER. 228S

-_OMPILER OPTIONS:

0001 C

0002 C

0003 C

0004 C

0005 C

0006 C

0007 C

0008

0009

0010

0011 C

0012 C

0013 C

0014

0015

0016 C

0017 C

0018 C

0019

0020

0021

0022

0023

0024

0025

0026 C

0027 C

0028 C

0029 C

0030 C

0031

0032 C

0033 C

0034 C

0035 C

0036

0037

0038.01

0039.01

0040.01

0041.01

0042.01

0043.01

0044.01

0045.01

0046.01

0047.02

0048.02

OO49.02

0050.02

0051.02

0052.02

0053.02

0054.01

0055.01 C

0056.01 C

0057.01 C

0058

0059

0060

0061

0062 C

0063 C

0064 C

0065

0066

<USER2>S343728217>STUDENT>DES I QN 1, F77

LISTINg INTL DCLVAR NOMAP CHECK NOBIg LOgL DYNM NOOFFSET LGO NOA_

NOFRN FPN NOLUNFREC NOSILENT NO_OPTIMISE NOIMPURE

AE 441: AEROSPACE DESIGN

GROUP D

PARAMETRIC SWEEP NUMBER 1

24 dAN 1990

***** INITIALIZE VALUES *****

REAL PI, RE, V, CBAR, RHO, VIS, KINVIS, PHI, g, RADIUS, N

PI=4.*ATAN(1. )

g=9.81

***** SET OUTPUT CODE *****

WRITE(lo*) 'ENTER OUTPUT CODE: "1" FOR SCREEN, "6" FOR PRINTER"

READ(I,*) IWR

***** INPUT PRELIMINARY DATA *****

WRITE(I,*) •ENSURE THAT ALL QUANTITIES ARE IN METRIC UNITS!!!"

WRITE(I,*) •ENTER TARGET REYNOLDS NUMBER"

READ(I,*) RE

WRITE(I,*) "ENTER LOCAL DENSITY"

READ(I,*) RHO

WRITE(I,*) "ENTER LOCAL FLUID VISCOSITY'

READ(I,*) VIS

WRITE(I,*) •ENTER DESIRED TURN RADIUS'

READ(I,*) RADIUS

***** PRELIMINARY CALCULATIONS *****

KINVIS=VIS/RHO

***** LOOP THE MEAN CHORD VALUES AND SOLVE THE REMAININg *****

***** EGUATIONS FOR THE OTHER QUANTITIES *****

WRITE(IWR,*) 'REYNOLDS NUMBER =',RE

DO 20 D=O., 12.

RADIUS=6. +D

WRITE(IWR,*)

WRITE(IWR,*)

WRITE(IWR,*)

WRITE(IWR,*)

WRITE(IWR, 85)

WRITE(IWRJ87)

WRITE(IWR,*) • "

DO 10 C=0.,38.

CBAR=.02+(C*.01)

V=(RE*KINVIS)/CBAR

PHI=ATAN(V**2. /(g-RADIUS))

N=I./COS(PHI)

PHI=PHI*180. /PI

WRITE(IWR, 86) CBAR, V, PHI,N

10 CONTINUE

20 CONTINUE

I •

'RADIUS =',RADIUS, '[M]"

¢ •

***** FORMATS *****

85 FORMAT(SX, 'MEAN CHORD', 5X, "VELOCITY', 5X, "BANK ANGLE', 5X, 'LOAD FAC

+OR " )

86 FORMAT(7X, F7. 3, 6X, F8. 3, 7X, F7. 3J 8X, F7. 3)

87 FORMAT(SX, "[M]', 11X, '[M/S] ',8X, "[DEg]')

***** STOP AND END *****

STOP

END

END OF COMPILATION CLOCKED .372 SECONDS
C Stephen T. Valentine (Design Qroup D)

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(

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(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

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1) C

2) C

3) C

4) C

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S) C

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35) C

36)

Stephen T. Valentine (Design group D)

AE 441: AEROSPACE SYSTEMS DESIGN

_ritten 02-15-90

LIFTING-LINE WINe ANALYSIS (Program B)

LIBRARY 'PLOTLIB'

...******************* vat lab led • c lira t t one *****************

REAL THETA (8), THETAP (20), ALPHA ( 11 ), ALPHAa ( 11 ), COEFM (8, S )

REAL 80LN(8, 11),ALPHMU(S, 11),CL(ll),CDI(11),Y(20, 1)

REAL gAMMA(20, 11),LIFT(20, 11),DRAO(20, 11),CLT(20, 11),CDT(20, 11

REAL COO(20, t I ), L, LCS, MU

INTEQER WK1 ( S ), WK2 (S), NP ( 11 )

• *********************** in i t i• 1 inp u t ***********************

WRITE (1,*) "enter output port number:"

READ (1,*) IWR

WRITE (1, e)

WRITE (I,*) "enter air•oil zero-lift angle ot/ attack'[deg]"

READ (1,*) AOL

• ******************* tnt t i•l output **************************

WRITE (IWR,*) " LIFT DISTRIBUTION DATA"

WRITE (IWR,*)

WRITE (IWR, 1)

FORMAT (6X, 'ALPHA', 14X, "CL', 12X, "CDI', 12X, 'Li_t')

WRITE (IWR, 4)

FORMAT (" £deg] ",21X, "[Ib]')

IOPT--111

constants declarations ***************

37) C ***** freestream *****

38) DENSTYm2.3769E-3

39) VELCTY-23.0

40) Gin_mO. 5*DENSTY*(VELCTY**2)

41) C **i** _ing . 

42 ) CHORD_K).6833333

43) C SPAN-8.0

C ARe11.72

C WAREA"SPAN** 2/AR

WAREAB5.5

5PANBWAREA/CHORD

C ***** alr_oll *****

C LCS=5.02

LCS=5. 5

AOL=-4. 0

44)

45)

46)

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49 )

50 )

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74)

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10

C

20

C

PIB4.*ATAN(1. )

DEQRADBPI/180.

MU=CHORD*LCS/(4.*SPAN)

AOI.=AOL,*DE_RAD

DO I0 I=1,8

THETA(I)=(90./8.)*REAL(I)*DEGRAD

CONTINUE

ALPHA(1)=6.0*DEQRAD

ALPHAa(1)=ALPHA(1)-AOL

NP(1)"'JO

DO 20 J'1.10

NP(J+I)=20

ALPHA(J+I)=DEQRAD*(ALPHA(J)/DEQRAD+0.2)

ALPHAa(J+I)=ALPHA(J+I)-AOL

CONTINUE

********* determine matrix o_ Fourier coefficients **********_

DO 40 I=l,S
C Stephen T. Valentine (Design Group D)

(

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2

3

DOL_._" *REAL(J)-1. J=1,8

COEFM(I,J)-(SIN(L.THETA(I)))*(1.+L*MU/SIN(THETA(I)))

CONTINUE

CONTINUE

DO 60 dm1,11

DO DO 1"1,8

ALPHMU(I, J)m(ALPHAa(J))*MU

CONTINUE

CONTINUE

CALL INVRT (COEFM, 8, 8, WK1, WK2)

CALL MULT (COEFM, ALPHMU, SOLN, 8, 8, 11, 8, 8, 8)

DO 80 I=1,11

CL(I)=PI*AR*SOLN(1, Z)

SUM=I.0

DR 70 J=2,8

L_'J.*REAL(J)-I.

SUMISUM+L*(SDLN(J,I)/SOLN(1, I))**2

CONTINUE

CDI(I)=((CL(I)**2)I(PI*AR))*SUM

E=I./SUM

CONTINUE

***************** calculate circulation distribution ********

DR 110 I-1,20

THETAP(I)=REAL(I)*(90./20. )*DEGRAD

Y(I,1)=COS(THETAP(I))

OR 100 d=1,11

SUM=O.O

DR 90 KI1,8

L=2.*REAL(K)-I.

SUM=SUM+SOLN(Ke d)*SIN(L*THETAP(I))

OAMMA(Ied)=2.0*SPANeVELCTY*SUM

LIFT(I,J)=OAMMA(IeJ)*DENSTY*VELCTY

CLT(I,J)-LIFT(I,J)/(O. 5*DENSTY*(VELCTY**2)*CHORD)

CDO(I,J)=1.7237E-2-(1.698E-2)*CLT(I,J)+(O. 09021)*(CLT(I,J

CDO(I,J)iCDO(I,J)-(O. 15822)*(CLT(Ied)**3)

CDO(IeJ)lCDO(I,J)+(9.0225E-a)*(CLT(IeJ)**4)

CDT(I,J)=CDO(I,J)+(CLT(I,J)**2)/(PI*E*AR)

DRAO(IeJ)=O. 5*DENSTY*(VELCTY**2)*WAREA*CDT(I,d)

CONTINUE

CONTINUE

CONTINUE

DO 120 I=le 11

WRITE (IWR,*)

CONTINUE

ALPHA(I)/DEORADeCL(I)eCDI(I),Gtnf*WAREA*CL(I)

WRITE (IWRe *)

WRITE (IWRe *)

WRITE (IWRe *)

WRITE (IWR,*)

WRITE (IWR,*)

DO 140 I_1,11

WRITE (IWRe*)

WRITE (IWRe*)

WRITE (IWRe*)

WRITE (IWR, 2)

WRITE (IWRe*) '

WRITE (IWR,*)

DO 130 d-le20

WRITE (IWR, 3)

CONTINUE

CONTINUE

'Span Efficiency Factor is ',E

'Spanwise Lift Distribution:'

"Incidence-', ALPHA(I )/DEGRAD, " deg1"ees "

£1blft] [Ib] '

Y(J, 1)e LIFT(de I), DRAG(d, I ), CLT(d, I ), CDT(d, ]

FORMAT

+

FORMAT

(2X, 'Y/(b/2)',7X, 'LIFT/unit span',SX,

'DRAG', 9X, 'CLT', 12X, 'CDT')

(2X, F7. 5, 8X, F7. 5, 8X, F7. 5, 8X, F7. 5, SX0 F7. 5)
c Z-3

Stephen T. Valentine (Design Q_oup D)

153)

154) C CALL TPLOT ( IOPT, Y, CLT, 20, 20, 5)

155) C CALL TLABEL ('Yl(b/2)', 'Lt_t Coefficient')

156) C CALL TITLE ('SPANWISE LOAD DISTRIBUTION (aOL= -4 degrees)') 157)

1_8) C PAUSE

159)

160) C CALL TPLOT (ZOPT, Y, CDT, 20,20,5)

161) C CALL TLABEL ('Y/(b/2)', "D_ag Coefficient')

162) C CALL TITLE ('SPANNISE DRAQ DISTRIBUTION (aOLm

163)

164) STOP

165) END

-4 degrees)')
1) C

2) C

3) C

4) C

5) C

6) C

7) C

8) C

9) C

10) C

11) C

12) C

13) C

14) C

15) C

16) C

17) C

18) C

19) C

CLOW=LIFT COEFFICIENT AT ZERO ANGLE OF ATTACK (WINg)

CLAW=LIFT CURVE SLOPE El/tad] (WINg)

CMACW=MOMENT COEFFICIENT ABOUT THE AERODYNAMIC CENTER (WINg)

XACW=X-POSITION OF THE AERODYNAMIC CENTER [_t] (WINg)

{mesured _rom the origin o_ the coordinate sgstem}

XACCB=AERODYNAMIC CENTER DIVIDED BY MEAN CHORD (WINg)

IW=INCIDENCE ANGLE [deg] (WINg)

S=PLANFORM AREA [_t**_] (WINg)

ARW=ASPECT RATO (WINg)

CBAR=MEAN CHORD [@t] (WINg)

CMOW=ZERO LIFT MOMENT COEFFICIENT (WINg)

CMAW=MOMENT CURVE SLOPE C1/rad] (WINg)

20) C .................

21) C

22) C

_3) C

24) C

25) C

26) C

27) C

28) C

29) C

30) C

31) C

32) C

33) CIT=INCIDENCE ANGLE [deg]

ST=PLANFORM AREA [@t**2]

BT=SPAN [_t]

LT=DISTANCE FROM TAIL A.C. TO CENTER OF gRAVITY

ART=ASPECT RATIO

CBART=MEAN CHORD

XACT=X-POSITION OF THE AERODYNAMIC CENTER [£t]

{measured @tom the origin o@ the coordinate

CLOT=LIFT COEFFICIEN_ AT ZERO ANGLE OF ATTACK

CLAT=LIFT CURVE SLOPE [I/_ad]

CMOT=ZERO LIFT MOMENT COEFFICIENT

CMAT=MOMENT CURVE SLOPE El/tad]

(TAIL)

(TAIL)

(TAIL)

(TAIL)

(TAIL)

(TAIL)

(TAIL)

sgstem}

(TAIL)

(TAIL)

(TAIL)

(TAIL)

34) C

35) C

36) CCMOF=ZERO LIFT MOMEN_ COEFFICIENT

CMAF=MOMENT CURVE SLOPE El/tad]

(FUSELAGE)

(FUSELAGE)

37) C

38) C

39) C

40) C

41) C

42) C

43) C

44) C

45) C

46) C

47) C

48) C

49) C

50) C

51) C

52) C

53) C

54) C

55) C

56) C

57) C

58) C

59) C

60) C

61) C

62) C

63) C

64) CXCOCB=CENTER OF gRAVITY POSITION DIVIDED BY MEAN CHORD

{measured _rom the leading edge o_ the _ing}

RHO=FREESTREAM DENSITY

VEL=FREESTREAM VELOCITY

G=FREESTREAM DYNAMIC PRESSURE

E=DOWNWASH ANGLE

EO=DOWNWASH ANgLE'AT ZERO ANGLE OF ATTACK

DEO=CHANgE IN DOWNWASH ANGLE WITH ANGLE OF ATTACK

CMO=MOMENT COEFFICIENT AT ZERO ANGLE OF ATTACK

CMA=CHANgE IN MOMENT COEFFICIENT WITH ANGLE OF ATTACK

SFNP=STICK FIXED NEUTRAL POINT/MEAN CHORD

{measured _rom the leading edge o_ the _ing}

SFSM=STICK FIXED STATIC MARGIN

{as a _actlon o# the mean chord}

ALFA=TRIM ANGLE [tad]

ALFD=TRIM ANGLE Eden]

CLW=WINg LIFT COEFFICIENT

CLT=TAIL LIFT COEFFICIENT

CLTOT= TOTAL AIRCRAFT LIFT COEFFICIENT

LW=WINg LIFT gENERATED

FT=TAIL LIFT gENERATED

LTOT=TOTAL LIFT gENERATED

CDO=FORM DRAg COEFFICIENT

CDTOT=TOTAL DRAg COEFFICIENT

DTOT=TOTAL DRAg gENERATED

LODTOT=TOTAL LIFT TO DRAg RATIO

65) C

66)

67)

68)

69)

70)

71)

72) C ....

REAL IW, IT(21),CMOW(21,21),VH(21,21),CMOT(21,21),ALFA(21,21)

REAL CMAW(21,21),CMAT(21,21),CMA(21,21),ALFD(21°21),CMO(21, 21)

REAL EO, DEO, ST(21),BT(21),ART(21),CLAT(21)_LT(21,21),CLOT

REAL CLW(21, 21), CLT(21, 21), LW(21, 21), FT(21_ 21), LTOT(21, 21)

REAL E(21,21),CDO(21, ;21),CLTOT(21,21),CDTOT(21,21),DTOT(21,21)

REAL LODTOT(21,21), XCOCB(21),SFNP(21),SFSM(21)

73) C

74) CINITIALIZE "HARD" VALUES

75) CLOW=,2596

7_) CLAW=3. 7!99
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149)

150)

151)

152)

CMACW=-O. 093

XACCB=.25

S=5.46

ARW=11.72

CBAR=.68333

XACT=2.935

CLOT=OoO

CBART=.4792

CMOF=-O. O002

CMAF=O. 0341

RH0=.0023769

PI=4.*ATAN(I. )

CONTINUE

WRITE(I,*)

READ(I,*)

WRITE(I,*)

READ(I,*)

WRITE(I,*)

WRITE(I,*)

READ(I,*)

WRITE(I,*)

READ(I,*)

G=.5*RHO*VEL**2

'ENTER OUTPUT CODE: El] FOR SCREEN, [6] FOR PRINTER'

IWR

"ENTER WINg INCIDENCE ANGLE [DEGREES]"

IW

'ENTER X-POSITION OF AERODYNAMIC CENTER OF WINg [FT]"

'{measured _rom the origin o@ the coordinate sgstem}"

XACW

'ENTER FREESTREAM VELOCITY [FT/S]'

VEL

C

C

NRITE(IWR,95) IW

WRITE(IWR, g6) XACW

WRITE(IWR,97) XACT

WRITE(INR, 99) VEL

DO I0 1=1,6

IT(1)=-I.5+REAL(I-I)*. 5

WRITE(IWR,*) ' '

WRITE(IWR,*) ' '

WRITE(IWR, 91) IT(I)

WRITE(IWR, 92)

WRITE(IWR,*) " '

DO 20 d=I,21

XCgCB(J)=.3333

ST(d)=.4+REAL(d-1)*.05

BT(d)=ST(d)/CBART

ART(J)=BT(d)**2./ST(J)

CLAT(d)=(2.*PI)/(I.+(2. /ART(d)))

C

C .......... O' VALUES FOR THE MOMENT COEFFICIENT

C

CMOW(V, I)=CMACW+CLOW*(XCQCB(J)-XACCB)

LT(J, I)=(XACT-XACW)+(XACCB-XCQCB(J))*CBAR

VH(J, I)=(ST(d)*LT(J, I))/(S.CBAR)

EO=((2.*CLOW)/(PI*ARW))*180. /PI

DEO=(2.*CLAW)/(PI*ARW)

CMOT(d0 I)=VH(d, I)*CLAT(d)*(EO+IW-IT(I))*PI/180.

CMO(d, I)=CMOW(d, I)+CMOT(d, I )+CMOF

C

C ........ A" VALUES FOR THE MOMENT COEFFICIENT

C

CMAW(J, I )=CLAW*(XCgCB(d)-XACCB)

CMAT(J, I )=-VH(d, I )*CLAT(J)*( 1. -DEO)

CMA(J, I )=CMAW(d, I )+CMAT(J, I )+CMAF

C

C .......... STICK FIXED NEUTRAL POINT

C

SFNP(d)=XACCB-(CMAF/CLAW)+VH(d,I)*(CLAT(d)/CLAW)*(1.-DEO)

SFSM(d)=SFNP(d)-XCQCB(J)

C

C FINDINg ANGLE OF ATTACK TO TRIM AND LIFT FORCES

C

ALFA(d, I)=CMO(V, I)/(-CMA(V, I))

ALFD(J, I )=ALFA(J, I )-180. /PI

CLW(V, I )=CLOW+CLAW*(ALFA (d, I )+( IW*PI/180. ) )

E(J, I)=(2.*CLW(J,I))/(PI*ARW)

CLT(J, I)=CLOT+CLAT(J)*(ALFA(J, I)-(IW*PI/180. )-E(J,I)

+ +(IT(I)*PI/180. ))

LW(V, I)=CLW(J, I)*G*S

FT(J, I)=CLT(J, I)*G*ST(d)
153)

154)

155) 156) C

157) C--

158) C

159)

160)

161)

162)

lb3)

1&4) C

165) C166) C

167)

168)

169) C

170)

171)

172) C

173) C--

174) C

175)

176)

177)

178)

179)

180)

181)

1821

183)

1841

185)

186)

187)

188)

189)

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192)

193)

194)

195)

1961

197)

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202)

203)

204)

2051

206)

207)

208)

209)

210)

211)

212)

213)

214)

215)

216) C

217) C--

218) C

219)

220)

221)

222)

223)

224)

2251

2_6)

2_7)

228) C

LTOT(J, I )=LW(J, I )+FT(J, I )

LTOT(,/, I )=LTOT(J, I )*. 966

CLTOT(d, I)=LTOT(d, I)/(G*S)

CALCULATINg DRAg FORCES

CDO(d, I)=(. 14885+(ST(J)*.O08))/S

CDO(d, I)=1. 15*CDO(J, I)

CDTOT(J,I)=CDO(d, I)+(CLTOT(J, I)**2/(.b815*PI*ARW))

DTOT(J, I)=CDTOT(J, I)*G*S

LODTOT(J,I)=CLTOT(J, I)/CDTOT(J,I)

PRINTINg STATEMENTS

WRITE(IWR, 90) ALFD(d, I),LTOT(d, I),VH(d, I)

+,LODTOT(d, I),SFNP(J), SFSM(d),CMA(J, I),ST(J)

20 CONTINUE

10 CONTINUE

CONTINUE

WRITE(I,*)

+/D RATIO,

+POINT, [7]

2

PLOTTINg STATEMENTS

'ENTER PLOTTINg OPTION: [1] = LIFT, [2] : DRAg, [3] : L

[4] - CM)al_a, [5] = TRIM ANGLE OF ATTACK, [6] = NEUTRAL

= STATIC MARGIN'

READ(I,*) NUM

IF(NUM. EG. I) THEN

CALL TPLOT(-O11,ST, LTOT, 21,21,6)

CALL TITLE('TOTAL LIFT gEN£RATED [Ib] vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [@t**2]', 'TOTAL LIFT gENERATED lib]')

ELSE IF(NUM. EG. 2) THEN

CALL TPLOT(-O11,ST, DTOT, 21,a1,&I

CALL TITLE('TOTAL DRAg gENERATED [ib] vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [@t**2]', 'TOTAL DRAg gENERATED lib]')

ELSE IF(NUM. EG. 3) THEN

CALL TPLOT(-O11, ST, LODTOT, 21,21,6)

CALL TITLE('LIFT TO DRAg RATIO vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [@t**2]', 'LIFT TO DRAg RATIO')

ELSE IF(NUM. EG. 4) THEN

CALL TPLOT(-O11,ST, CMA, 21,21, 1)

CALL TITLE('Cm)aI@a vs. TAIL AREA [Ft**2]')

CALL TLABEL('TAIL AREA [Ft**2]', 'CM)al£a')

ELSE IF(NUM. EG. 5) THEN

CALL TPLOT(-O11,ST, ALFD, 21,21,_)

CALL TITLE('TRIM ANGLE OF ATTACK [deg] vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [@t**2]', 'TRIM ANGLE OF ATTACK [deg]')

ELSE IF(NUM. EG.&) THEN

CALL TPLOT(-O11,ST, SFNP, 21,21,1)

CALL TITLE('NEUTRAL POINT vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [Ft**2]', 'NEUTRAL POINT')

ELSE IF(NUM. EG. 7) THEN

CALL TPLOT(-O11,ST, SFSM, 21,21, 1)

CALL TITLE('STATIC MARGIN vs. TAIL AREA [@t**2]')

CALL TLABEL('TAIL AREA [@t**2]', "STATIC MARGIN')

ELSE

ENDIF

WRITE(I,*) "MORE PLOTS ? £1] FOR YES, [NOT 1] FOR NO"

READ(I,*) NUMMM

IF(NUMMM. EG. 1) @OTO 2

WRITE(I,*) 'CONTINUE ? [1] YES, [NOT 1] FOR NO'

READ(I,*) NUMM

IF(NUMM. EG. 1) @OTO 1

FOR MAT STATEMENTS

90 FORMAT(2X, F7. 4, 2X, F6. 4, 2X, F&. 4, 2X, F8. 4, _X, F8. 4, _X, F8. 4, 2X, FS. 4, 2X,

+F8. 4)

91 FORMAT(2X, 'TAIL INCIDENCE ANGLE IS ",F8. 4, " [DEGREES]')

92 FORMAT(4X, "ALFA',4X, "LIFT',3X, "VolRat',SX, 'LoD)t', 5X, 'SFNP',

+6X, "SFSM',SX, "CM)al@a',4x, 'Stail')

95 FORMAT(aX, 'WIN@ INCIDENCE ANGLE : ',F8.4)

96 FORMAT(2X, 'WIN@ AERODYNAMIC CENTER : ',F8.41

97 FORMAT(2X, 'TAIL AERODYNAMIC CENTER = ',F8.41

99 FORMAT(2X, 'FREESTREAM VELOCITY [FT/S] = ',F8.4)
ITEM

(

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(

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(

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(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

C

(

(

(

(

(

(

C

(

(

C

C

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

(

1) C

2) C

3) C

4) C

5) C

6)

7)

8)

9) C

10) C

11)

12)

13)

14)

15)

16)

17)

18) C

19) C

20) C

21) C

22) C

23) C

24) C

25) C

26) C

27)

28)

29)

30)

31) C

32) C

33)

34)

35)

36)

37) C

38) C

39) C

40)

41)

42)

43) C

44) C

45)

46) C

47) C

48)

49) C

50) C

51) C

52) C

53)

54)

55) 20

56) 10

57) C

58) C

59)

bO)

62)

63)

64)

65)

b6) 40

b7) 30

68) C

bg) C

70) C

71) C

72) C

73)

74)

SIZINg THE VERTICAL TAIL & RUDDER

gROUP D -- dEFFREY dANICIK

____.______

REAL SV(110),LV(110),BETA(110,110)

WRITE(I,*) 'ENTER INR: '

READ(I,*) IWR

******** FIXED PARAMETERS *********

PI=4.*ATAN(I. )

SN=5.46

B=8. O

ARW=11.72

ETAV=I.0

CLAV=(2.*PI)/2.45

CNBWF=-.O000721

******* VARYINg PARAMETERS ********

-- CONSTRAINTS SET BY gROUP DESIGN ANO "RULES OF THUMB"

Iv = 1.5 @t - 2.0 £t

Sv = .38 _t_'2 - .bO _t"2

tao = .62 - .72 Figure 2.20 Nelson

delta rudder max = 25 degrees

***********************************

WRITE(I,*) 'ENTER VALUE OF TAO " '

READ(I,*) TAO

WRITE(I,*) 'ENTER VALUE OF DELTA RUDDER " '

READ(I,*) DELTAR

DO 10 I=37,42,1

SV(I)=REAL(1)/IO0.

DO 20 d=18,20, 1

LV(J)=REAL(_)/IO.

***** CALCULATINg CNB OF VERTICAL TAIL --eqn. 2.80 NELSON *****

--- SDEF = SIDEWASH & TAIL EFFICIENCY FACTOR (EMPIRICAL FORMUL_

VV=SV(I)*LV(J)/(SN*B)

SDEF=.724+3. 06*((SV(1)/SW)/2. )+0.2+O. O09*ARN

CNBV=VV*CLAV*SDEF*PI/180.

***** CALCULATINg CNB OF ENTIRE AIRPLANE *****

CNB=CNBWF+CNBV

***** CALCULATINg CN-DELTA-R -'e_n.s 2.86 & 2.87 NELSON *****

CNDELR=-ETAV*VV*CLAV*TAO*PI/180.

***** CALCULATINg BETA (THE HEADINg ANGLE) *******

-- THIS IS DONE BY SETTINg CN = 0 FOR TRIM CONDITION IN

EGUATION CN = CNB*BEIA + CNDELR*DELTAR

BETA (d, I )=-C NDELR*DELTAR/CNB

WRITE( IWR, *) CNB, CNDELR

CONTINUE

CONTINUE

WRITE(INR,*) 'TAO = ",TAO, 'DELTA RUDDER =',DELTAR

WRITE(IWR,*) ' "

NRITE(IWR,*) " Sv (_t^2) ', " Lv (_t) ", "

DO 30 I=37, 42, 1

DO 40 d=18, 20, 1

WRITE(IWR,*) SV(I),LV(d),BETA(d, I)

CONTINUE

CONTINUE

CALL TPLOT(-O110SV_BETA_100_100,6)

CALL TITLE('BETA vs. Sv')

CALL TLA3EL('Sv (_t'_2)', 'BETA (degrees)')

STOP

END

BETA (DEGREE
C

I) C

2) C

3) C

4) C

5)

6)

7)

8) C

9) C

10)

11) C

12) C

13) C

14) C

15)

16)

17)

18)

19)

20) C

21) 20

22) 10

23) C

24) C

25)

26)

27)

28) C

29)

30)

**************************************

PROGRAM TO GENERATE Cl vs BETA GRAPH

-- GROUP D dEFFREY L. dANICIK

REAL BETA(40),CL(40,40),GAMMA(40)

WRITE(I,*) "INPUT IWR: "

READ(I,*) IWR

**** ASPECT RATIO EFFECT ON CLB FOUND BY FIGURE 3.9

CLBGAM=-.O0027

  VARYING PARAMETER --DIHEDRAL ** ***

..... RULE OF THUMB" GAMMA = 5 DEGREES - 8 DEGREES

DO 10 I=1, 10

GAMMA(I)=6.+REAL(1)

DO 20 d=9,28,1

BETA(J)=REAL(d)

CL(J, I)=CLBGAM*GAMMA(I)*BETA(J)

WRITE(IWR,*) gAMMA(I),BETA(J),CL(J,I)

CONTINUE

CONTINUE

****** PLOTTING ROUTINE *******

CALL TPLOT(-O11,BETA, CL, 40,40, 10)

CALL TITLE('CI vs. BETA')

CALL TLABEL('BETA (degrees)', "ROLL MOMENT COEFFICIENT Cl')

STOP

END
EDWARD DIESER

I) C

2) C

3) C

4) C

5)

b)

7)

8)

9)

10) C

II) C

12) C

13)

14)

15)

16)

17)

18)

19)

20)

21) C

22) C

23)

24)

26)

27)

28)

30)

31) C

32)

33)

34)

35)

36)

37)

38)

39)

40)

EDWARD DIESER

AERO DESIGN

TRADE STUDY

22 MARCH 1990

REAL IC(15),dC(25),DIHED(15),YAWA(_5),DAL(250 15)

PI=4.*ATAN(1. )

YAWA= YAW ANGLE

DAL= DELTA ALPHA

DIHED=DIHEDRAL ANGLE

WRITE(I,*) 'ENTER IWR (1=TERMINAL, 6=PRINTER)'

READ(I,*) IWR

DO 10 I=1015

IC(1)=6.+REAL(I)

DIHED(I)=IC(I)*PI/180.

WRITE(IWR,*) "DIHEDRAL ANGLE=', It(I) ,'DEGREES"

WRITE(IWR,*) "YAW ANCLE (DEGREES) DELTA ALPHA (DEGREES)"

DO 15 d=1,25

OC(d)=REAL(d)

YAWA(d)=UC(d)*PI/180.

DAL(d, I)=ATAN(SIN(YAWA(d))*TAN(DIHED(I)))

DAL(d0 I)=DAL(O° I)*180./PI

15 CONTINUE

15 WRITE(IWR,*)" ",OC(O), " ',DAL(O, I)

10 CONTINUE

+L (deg) ")

CALL TLABEL('YAW ANOLE

STOP

END

CALL TPLOT(-O11,dC, DAL, 25,25,15)

CALL TITLE('ANOLE OF ATTACK vs. YAW ANGLE _or various WINg

(degrees)', 'DELTA ALPHA (degrees)')

DIHEDRA
EDWARD DIESER

1) C

2) C

3) C

4) C

5)

6)

7)

8)

9)

10)

11)

12)

13)

14)

15)

16) C

17) C

18)

19)

20)

21)

22)

23)

24)

25)

26)

27)

28) C

29) C

30)

31)

32)

33)

34)

35)

36)

37)

38)

39) C

40) C

41)

42)

43) C

44) C

45) C

46)

47)

48)

49)

50)

51)

52)

53)

5

7

30

_0

20

10

10

EDWARD DIESER

AERO DESIGN

TRADE STUDY

22 MARCH 1990

REAL ROLMO(15),ROLI(15),ALPHA(15),LI(15),LIP(15),RIP(15),CL(15)

G=.5-.00237-23.*-2

S=4./12.*.683

PI=4*ATAN(I. )

PC=O.

WRITE(I,*) "ENTER IWR (1=TERMINAL. 6=PRINTER)"

READ (1,*) IWR

DO 5 R=1,11

ROLMO(K)=O.O

ROLMO(1)=O.O

LI(2)=2

DO 7 K=3,13

LI(K)=2+REAL(K-2)*4.

DO 10 I=1,15

ALPHA(I)=&.+REAL(I)

CL(I)=.6+ . 1*REAL(I)

WRITE(IWR,*) "ANGLE OF ATTACK=',ALPHA(I),

WRITE(IWR,*) 'PANEL SECTION # LENGTH

PC=O.

DO 20 d=2,13

PC=PC+I

ROLI(J)=CL(I)*G*S*LI(J)

ROLMO(J)=ROLMO(J-1)+ROLI(d)

LIP(J)=LI(J)/48.

RIP(J)=ROLI(J)

DO 30 K=1,11

RIP(J)=RIP(J)/ROLMO(11)

CONTINUE

WRITE(IWR,*) " ",PC," ',LI(J)," ",ROLI(J)

WRITE(IWR,*) "ROLL MOMENT=', ROLMO(12)

NRITE(IWR,*) '

CONTINUE

"(deg)'

SECT. ROLL MOMENT"

STOP

END

°CALL TPLOT(-O11,ALPHA, ROLMO, 15, 15, 1)

CALL TITLE( 'ROLL MOMENT vs. ANGLE OF ATTACK')

CALL TLABEL('ALPHA (deg?ees)', "ROLL MOMENT (ft.-lb.)")
 TEm

TORNADO 1 lil--6 PROF'ELLER

Performance Estimate

Analysis_, o_p,:-' '- ; ade _.:_n=.,tr,e'.ir_

_..r.-, _,:oL:RIPTION'2 Blades; IZlIn, Dia,," '>='_ : "_- " P_a _ '-:et_t _' " . ' . ......

L,_H=CONDITIONS: Calculahons adiust-:_ •-_rr:eit'er_a-h ...... at '=iJ ...._ "_ '

Blade Measurements

Thrust_ Power_ Efficiency, and Velocities

Angles of Attach: 'n--,._,--}._,_,_e_

x ,J, _' .2_ , .'_ , Ta ....:"v 4_. , 49 , 55

• ; ',_- i5.2 12,i _.

' 7,2 i,7 -*., a

t

'= :_...; [*..3 :'.3 " .'t 5.5 '; :' ,._

._ _a,3 8,$ 7 i: = '_ _,.i 3 ' =

7 _ 7 7 ,_.4 5.1 " ;. -._ ' "

"= _.a 7.2 8 4,'a " 5 Z,4 ; 3

.; 7.# 6.9 5.6 4. =_ _.,.._ 2,3 1,2

_= 7,5 6.4 %; 1.3 7.2 _" I I

,_ ¢,7 g,7 1.8 7.8 ,.,Q _,_ 1

._5

l -: ,_

•I -!

,! -,Q

,i

a -.;

.7 --T

-4, " -*

-_ ° -i.:'

--." -7.-

-" " -7, 3

Reynolds Number ':Mi!lion_!

Y J: .24 .29 ,34 ._ i' :; :_ " ,

,7 ._4 ._3 ,_3 ._2 ,_2. , _2 ,_2. . _,.'_ ,_2.... ,C _"

,6 ._ ._7 ,_6 ._5 ,_5 ,_ ._4 ._ ,_ ,_; .-U

,;5 ,_? ,_7 ,_ ,@_ ._5 ._4 .;_4 ,_ ,_ ._; ,__

,_ ,a_ ,e ,06 ,_5 _5 ._4 ,_,4 ._4 ._7 ,_7 ....

.';5 4_ ,_4 ._4 _.T 3' _7 e2 .@2 a, ,a_ '"

,, _,, e ,.c d ,_ata for _.,e,__ed _irg2i t SeCti3r.

ORIGINAL PAGE IS

OF POOR QUALITY
PROPELLER DATA SHEET

A)

B)

C_

Dl

E)

G)

H_

I)

Propeller Designation: TORNADO I_-6

Number of Blades: 2 Diameter: 10 (Inches)

Select one of the following airfoil sections:

1) INVISCID FLAT PLATE

2) THIN FLAT PLATE

3_ SYMMETRICAL

--> 4) CLARK Y

5) RAF-6

Blade thickness may be entered as either:

I_ Fraction of chord

--> 2) Inches

Blade data may be entered at radial locations specified as:

I) Fractional Radius

-- 2) Inches

F) Radius* at which blade setting is measured: 2.125

Blade setting (i.e. ref angle for whole blade): 24.2

Enter the number of radial data positions: (3-9): 9

Data Point Radius* Chord Thickness* Angle

I: .625 .625 .305 56.8

_" 1 I_= 7_ 225 40 _

3: 1.625 .8125 •191 36.4_

4: 2.125 .875 .165 24.2

5: 2.625 .9375 .147 19.99

6: 3.125 .9375 .134 16.9_

?:. 3.625 .81_5_ . 114 14. 76

8: 4.125 .6875 ._91 13-_53

9: 4.625 .5 ._58 11.67

J)

_'-:i)

i )

M)

N )

Select desired refinment of analysis:

--_ I) Analysis by simple blade element theory.

2) Analysis including induced velocity.

3) Analysis including induced velocity and tip losses.

These CI/Cd coefficient adjustments may be selected

--> I) No Cl/Cd adjustments

_) Math number adjustment

3) Reynolds number adjustment

4) Math and reynolds number adjustments

Select altitude in thousands of feet: .02

Specify one of the following:

--> I) Airspeed FIXED at: 15 MPH

2) Propeller RPM FIXED at:

_ange of Advance Ratio to be used in calculations:

J min: _ J max: 1

.°- , -,

ORIGINAL PAGE I$

OF POOR QUALITY
CASE - GROUPD

WGT = 3.000

SREF = 5. 460

RH0 = 2.3780E-03

CLTO = .3000

CDTO = 4.0000E-02

CLMAX = 1.200

SMAX = 300.0

MU = .2000

DIA = .8333

BVOLTS = 8. 400

KT = .5500

KV = 5.0000E-04

RARM = 6.0000E-02

RBAT = 4.7000E-02

FUSAMP = 20.00

GF.ABAT = 2.210

DT = 5.0000E-02

TMAX = 60.00

NJ = 10

.0000 .1040 7.0000E-02

.2400 9.2000E-02 5.0000E-02

.2900 8.5000E-02 7.9000E-02

.3400 7.9000E-02 4.7000E-02

.3900 7.2000E-02 4.5000E-02

.4900 5.4000E-02 3.9000E-02

.5500 4.3000E-02 3.5000E-02

.6500 2.0000E-02 2.1000E-02

.7500 -6.0000E-03 7.0000E-04

lO.0O .00oo -lOO.O

V TAKEOFF = 23.55

MAX CURRENT DRAW(amps) = 39.25

MAX MOTOR POWER(hp) = .1798

MAX MOTOR POWER(watts) = 134.1

STATIC THRUST (lb)= 1.186

STATIC CURRENT DRAW (amps)= 16.73

STATIC PROP RPS= 99.71

TATIC PROP RPS= 99.71ps)= 16.73) = .3840

V AT TO (FT/SEC) = 23.76

DISTANCE(FT) = 45.06

BATTERY DRAIN(mahs) = 16.03

ADVANCE RATIO AT TO = .2892

THRUST(LB) AT TO = .9306

LIFT(LB) AT TO(BEFORE ROTATION) = 1.080

DRAG(LB) AT TO(BEFORE ROTATION) = .1440

FRICTION(LB) AT TO(BEFORE ROTATION) = .3840

CURRENT DRAW AT TO (AMPS) = 18.03

• .¢ sP s - sX
t

GROUPD

3.0

5.46

.002378

.3

.04

1.2

300.

.2

.83333

8.4

.55

.0005

.06

.047

20.

2.21

.05

60.

I0

0.

.24

.29

.34

.39

.49

.55

.65

.75

i0.

.104

•092

.085

.079

.072

.054

.043

.02

-.006

0.

t WEIGHT IN LBS

t WING REF AREA IN FT2

! AIR DENSITY SLUG/FT3

t CL AT TAKEOFF ATTITUDE

t CD AT TAKEOFF ATTITUDE

I CLMAX -

t LIMIT ON TAKEOFF DIST FT

t FRICTION COEFFICIENT

t PROP DIAMETER IN FT

r BATTERY VOLTAGE

! KT IN in-oz/amp

! KV IN volts/rpm

! armature resistance

! battery resistance

! FUSE AMPS - MAX DRAW

! gear ratio

t INTEGRATION TIME INCREMENT

' limit on take-off time (SEC)

! number of prop data points

.07

.05

.079

.047

.045

.039

.035

•021

•0007

-100.
_FORD UNIVERSITY FTN77 VER.

MP ILER

0001

0002

0003

0004

0005

0006

0007

0008

0009

0010

0011

0012

0013

0014

0015

0016

0017

O01S

0019

0020

0021

0022

0023

0024

0025.01

0026.01

0027.01

0028

0029.01

0030.01

0031.01

0032.01

0033.01

0034

0035

0036

0037

0038

0039

0040.01

0041.01

0042.01

0043.01

0044.0I

0045

0046.01

0047.01

0048.01

0049.01

0050

0051

0052

0053

0054.01

0055.01

0056

0057

0058

0059

228S

OPTIONS: LISTIN_ INTL NOMAP CHECK NO_IQ LOlL DYNM NOOFFSET

FPN NOLUNFREC NOSILENT NO_OPTIMISE NOIMPURE

***************************************

C* AE446 - AEROSPACE STRUCTURES *

C* Compute_ program to determine *

C* the internal load distribution *

C* on an aircraft winq. *

C* Written by Michael H. Pa_k *

***************************************

PROCRAM STRUCI

REAL

+

ITEI_ IO

£USER2],S310768966}STUDENT]_STRUC I. F77

LgO NOANSI

X(50),CL(50),CD(50),CM(50),W(50),PY(50),PZ(50),DMX(50),

VY(50) ,VZ (50), MX(50), MY(50), MZ(50) ,C(50)

5

10

2O

30

90

I

20: 1:

NODEI

OPEN(UNIT=99, FILE='WINgl. DATA',STATUS='OLD ")

READ(99,*)RO, ALPHA, V, CR, CT, N

READ(eg,*)(X(I),CL(1),CD(1),CM(1),W(I), I=I,N)

CLOSE(UNIT=99)

G=O. 5*RO*V**2

PI=4.0*ATAN(1.0)

A=ALPHA*PI/180. 0

WRITE(1,_)'ENTER IWR: '

READ(I,*)IWR

DO 5 I=I,N

C ( I >=CR-((CR-CT)/X (N))*X (I )

CONT INUE

DO I0, I=I,N

PY(I)=G_C(1)_(CL(I)*COS(A)÷CD(I)*SIN(A))-W(1)*COS(A)

PZ(1)=Q*C(I)_(CL(I)*SIN(A)-CD(I)*COS(A))-W(1)*SIN(A)

DMX(I)=G*C(I)_*2*(-I.0*CM(I)-CL(1)*COS(A)/4.0-CD(1)_SIN(A)/4.0)

CONTINUE

VY(N)=O.O

VZ(N)=O.O

MX(N)=O.O

MY(N)=O.O

MZ(N)=O.O

DO 20 I=N-1, 1,-1

VY(I)=VY(I+I)+(PY(I+I)+PY(I))*(X(I÷I)-X(I))/2.0

VZ(I)=VZ(I+I)+(PZ(I+I)+PZ(1))*(X(I+I)-X(1))/2.0

MX(r)=MX(I+I)+(DMX(I+I)+DMX(1))*(X(I+I)-X(I))/2.0

CONTINUE

DO 30, I=N-1, i,-I

MZ(I)=MZ<I+I)+(VY(I+I)+VY(I))*(X(I+I)-X(1))/2.0

MY(1)=MY(I+I)+(VZ(I+i)+VZ(1))*(X(I÷I)-X(1))/2.0

CONTINUE

WRITE(IWR,*)' X Vg(Ib$) Vz Mx(@t-lbs) Mg

WRITE< IWR, *) '

Mz'

÷______ •

DO 90 I=I,N

WRITE( IWR, 1)X< I), VY(I), VZ(I), MX(1), MY(1), MZ(I)

CONTINUE

FORMAT(IX, FS. 2, IX, F9. 5, IX, Fg. 5, IX, Fg. 5, IX, F9. 5, IX, F9. 5)

STOP

END

3 OF COMPILATION CLOCKED • 572 SECONDS
Vw(lbs) Vz Mx(_t-lbs) M_ Mz

i0- _'.

O. O0

0.31

O. b3

O. 93

I. 24

1. 53

I. 82

2. 09

2.35

2, 60

2. 83

3. 04

3. 24

3.41

3. 56 3 70

3. 80

3. 89

3.95

3. 99

16.

14.

13.

12.

i0.

9.

8.

7.

5.

4.

3.

3.

2.

I.

I.

O.

O.

O.

O.

O.

23668 1.29189 -20.

83717 !.18189 -18.

44794 1.07263 -16.

07876 0.96484 -15.

74015 0.85929 -13.

44182 0.75669 -11.

19319 0.65776 -9.

00408 0.56321 -8.

88393 0.47376 -6.

84188 0.39013 -5.

88608 0.31299 -4.

02453 0.24303 -3.

26480 0.18096 -2.

61361 0.12748 -1.

07557 0.08321 -0.

65281 0.04858 -0.

34374 0,02372 -0.

14259 0.00831 O.

03600 0.00119 O.

00000 0.00000 O.

69218

81430

95127

11737

32781

59669

93779

36553

89396

53662

30560

21270

26930

48536

86707

41562

12537

01781

04280

00000

2. 41726

2.02908

1.67751

1.36366

1.08796

0.84989

0.64816

0.48087

0.34546

0.23893

0.15784

0.09857

0.05735

0.03044

0.01427

0.00561

0.00168

0.00031

0.0000_

0.00000

30,24091

25.36481

20.95403

17.02190

13.57305

10.59984

8,08474

6.00_41

4.31949

2.99679

1.99028

1.25361

0.73938

O. 40103

0,19466

0.08104

0.02686

0.00614

0.00066

0.00000
O. 00237, 7. 6, 2:3. O, 8. 2., 8. 2., _0

( 1) 0.00237,

( 2) 0.00000,

( 3) 0.31384,

( 4) 0.62572,

( 5) 0.93380,

( 6) 1.23608,

( 7) 1. 53072,

( 8) I. 8159&,

( 9) 2. 09000,

( 10) 2. 35116,

( 11) 2.59780,

( 12) 2.82844,

( 13) 3.04164,

( 145 3.23608,

( 15) 3. 41056,

( 165 3. 56404,

( 17) 3.69552,

( IS) 3.80424,

( 19) 3.88948,

( 20) 3.95076,

( 21) 3.98768,

7. 6,23. 0,8. 2,8. 2,_0

0 942.10, O. 04717, -0. 093, O. 375

0 94164, O. 04712, -0. 093, O. 375

0 94020, O. 04699, -0. 093, O. 375

0 93769, O. 04677, -0. 093, O. 375

0 9340&, O. 04644, -0. 093, O. 375

0 92920, O. 04602, -0. 093, O. 375

0 92279, O. 04547, -0. 093, O. 375

O. 91434, O. 04476, -0. 093, O. 375

0

0

0

0

0

0

0

0

0

0

O

0

90336, O. 04386, -0. 093, O. 375

88944, O. 04278, -0. 093, O. 375

87191, O. 04148, -0. 093, O. 375

84938, O. 03992, -0. 093, O. 375

_1984, O. 03803, -0. 093, O. 375

78121, O. 03582, -0. 093, O. 375

73152, O. 03332, -0. 093, O. 375

_802, O. 03057, -0. 093, O. 375

58601, O. 02754, -0. 093, O. 375

47931,0 02417,-0. 093,0. 375

34342, O. 02049, -0. 093, O. 375

I°015, O. 01727, -0. 093, O. 375
,-FORD

MP ILER

OO01

0002

0003

0004

0005

0006

0007

0008

O0O9

0010

0011

0012

0013

0014

0015

001b

0017

O01S

0019

0020

0021

0022

0023

OO24

0025

002_

0027

0028

0029

0030

003I

0032

0033

0034.0I

0035.01

0036.01

0037

0038.01

0039.01

0040.01

0041.01

0042.01

0043

0044

0045

0046

0047.01

00_8.01

004_.01

0050.01

0051.01

0052.01

0053.01

0054.01

0055.01

0056

0057.01

0058.01

0059.01

0060.01

0061.01

0062.01

0063

0064

0065

0066

0067

0068

0069

C070

0071

0072

0073

C074

0075

UNIVERSITY FTN77 VER. 228S _:USER2>S31076896b>STUDENTzSTITO'C'3_77 20:

OP T IONS : LISTINO INTL DCLVAR NOMAP CHEC_ NOBIO LOgL DYNM NOOFFSET LEO NOA!

NOFRN FPN NOLUNFREC NOSILENT NO_OPTIMISE NOIMPURE

***************************************

C* AE446 - AEROSPACE STRUCTURES *

C* Computar program to determine *

C* the centroid location and sec- *

C* tion properties relative to the *

C* centroid _or a nouni_orm air_oil *

C_ and the direct stress due to *

C_ axial, bending and thermal loads *

C* Written b_ michael H. Park *

***************************************

FROCRAM STRUC4

REAL E(50),A(50),Y(50),Z(50),PT(50),MZT(50),MYT(50), IYY, IZZ, IYZ

÷,SIg(50),ALPH(50),MY, MZ

CPEN(_9, FiLE='PROPSS. DATA,,STATUS='QLD ')

READ(99,_)N

REA_(99,_)(A(I), Z(I) Y(_),E(I),ALPH(I), I=I,N)

CLOSE(UNIT=99)

WRITE(I,_)'ENTER IWR '

READ(I,_)IWR

1 WRITE(t,*)'ENTER P, Mg, Mz, DT: "

READ(I,_)P, MY, MZ, DT

ATCT=O.O

YA=O.O

ZA=O.O

IYY=O.O

[ZZ=O.O

!YZ=O.O

I=!

EREF=E(i)

DO 5 I=I,N-I

IF(E(1).LT.E(I+I))EREF=E(1)

CCNTINUE

10

DO I0, I=I,N

ATOT=ATOT+(E(1)/EREF)*A(1)

YA=YA+(E(1)/EREF)_Y{I)_A(1)

ZA:ZA÷(E(1)/EREF)_Z(1)_A(1)

CONTINUE

YBAR=YA/ATOT

ZBAR=ZA/ATOT

DO 20, I=I,N

IYY=IYY+(E(1)/EREF)*A(1)*(Z(I)-ZBAR)_*2

IZZ=IZZ+(E(I)/EREF)*A(I)*(Y(1)-YBAR)**2

IYZ=IYZ+(E(I)/EREF)*A(1)*(Z(I)-ZBAR)e(Y(I)-YBAR)

2O

PT(I)=E(1)*ALPH(1)*DT

MYT(1)=E(1)_A(1)*Z(1)*DT

MZT(I)=E(I)_A(I)_Y(I)_DT

CONTINUE

3O

DO 30, I=I,N

B=E(I)*(P+PT(1))/(EREF*ATOT)

C=(E(1)/EREF)_((MZ-MZT(I))WIYY+(MY+MYT(1))*IYZ)*Y(1)

D=(E(1)/EREF)_((MY+MYT(1))*IZZ+(MZ-MZT(1))*IYZ)*Z(1)

SIO(1)=B÷(D-C)/(IYY_IZZ-IYZ**2)-E(I)*ALPH(1)*DT

CONTINUE

WRITE(!WR,_-) "YBAR = ', Y3AR, "in. "

WRITE(IWR,*) "ZBAR = ", Z3AR, "in. '

WRITE(IWR,*) "Igg = ', IYY, "in"4'

WRiTE(IWR,*) 'Izz = ', IZZ, "in"4"

WRII'E(IWR,*)'I_z = ", IYZ, 'in"4"

WRITE(IWR,*) "P ',P, 'ibs"

WP. ";RITE(IWR,_.) ITE(IWR,*) ;Mgz == ",',MZ, MY, 'in-lbs" 'in-lbs'

WR!TE(IWR,-_-) "DT = ", DT, 'd._grees F'

WRITE(IWR,*) ' "

WRITE( IWR, 98)

WRITE(IWR, _) .........
SAL,--QRDUNIVERSITY FTN77 VER. 228S

COMPILER OPTIONS:

C

C

C

C

CC

C

C

C

20

3O

4O

LISTINg INTL DCLVAR

NOFRNFPN NOLUNFREC

0001

0002

0003

0004

0005

0006

0007

0008

0009

0010

0011

0012

0013

0014

0015

0016

0017

0018

0019

0020

0021

0022

OO23

0024

OO25

0026

O027

0028

0029

0030

0031

0032

0033

0034

0035

0036

0037

0038

0039

0040

0041

0042

0043

0044

0045

0046

0047

0048

0049

0050

0051

0052

0053

0054

0055

0056.01

0057.01

0058.01

0059.01

0060

0061

0062

0063

0064.01

0065.01

0066

0067

0068

0069

0070.01

0071.01

0072.01

0073

0074

0075

-CUSER2>S343728217>STUDENT;'_

NOMAP CHECK NOBIg LOGL DYNM NOOFFSET

NOSILENT NO_OPTIMISE NOIMPURE

REAL IY, X(30),Z(30),W(30)

WRITE(l,*)

READ(1,_)

CONTINUE

XC=O.

ZC=O.

WT:O.

IY=O1

"ENTER OUTPUT

IWR

CODE: [I] FOR SCREEN, C6] FOR

X(1)=1. 125

Z(1)=2.625

W(1)=7.5

X(2)=4.0

Z(2)=2.625

W(2)=3.23

X(3)=6.0

Z(3)=0.625

W(3)=0.95

X(4)=6.0

Z(4)=2.28

W(4)=2.0

X(5)=18.94

Z(5)=2.44

W(5)=1.2

X(6)=9.0

Z(6)=3.0

W(6)=2.0

X(7)=14.834

Z(7)=2.25

W(7)=5.22

X(8)=12.62

Z(8)=0.75

W(8)=10.36

X(9)=33.67

Z(9)=4.25

W(9)=0.4574

X(10)=35. 58

Z(10)=4.0

W(10)=0.7226

X(11)=9.0

Z(li)=-3.0

W(11)=2.4

W(12)=5.46

WRITE(I,*)'ENTER X- AND

READ(I,_)X(12),Z(12)

N=I2

DO 20 d=l,N

X¢=XC+X(J)_W(J)

ZC=ZC+Z(J)*N(J)

_T=WT+W(J)

CONTINUE

XC_=XC/WT

ZC_=ZC/WT

Z-POSITIONS OF BATTERY PACK'

DO 30 K=I,N

IY=IY+((X(K)-XC¢)_*2÷(Z(K)-ZCG)_*2)_W(_)

CONTINUE

IY=IY/(144._16. _32.2)

WRITE(IWR, 90)

DO 40 L=I,N

{_RITE(IWR, 91)

CONTINUE

L,X(L),Z(L),W(L)

WRITE( IWR, *) " '

WRITE(IWR, 92) XCQ

WRITE(IWR, 931 ZCO

20:0

LgO N

PR INTER '
0076

0077

0078

0079

0080

0081

0082

0083

0084

0085

0086

0087

0088

0089

WRITE(IWR, 95)WT

_RITE(IWR, 94) IY

WRITE(I,*) 'CONTINUE ? [I] FOR YES"

READ(l, *) NN

iF(NN. EQ. 1) gOTO 1

90 FORMAT(2X, "COMPONENT ', 3X, "X-POS',4X, "Z-POS',2X, "WEIGHT [oz]')

91 FORMAT(SX, I2, 6X, F7. 4, 2X, F7. 4, 2X, F8. 4)

92 FORMAT(2X0 'X-POSITION OF THE C.Q. = ",F7. 4, 1X, "INCHES')

93 FORMAT(2X, 'Z-POSITION OF THE C.g. = ",F7.4, 1X, 'INCHES')

94 FORMAT(2X, "MOMENT ABOUT Y-AXIS {Iv> = ',F7.4, 1X, 'SLUg*FT**2")

95 FORMAT(2X, "TOTAL WEIGHT = ",F7. 4, 1X, 'OZ')

STOP

END

END OF COMPILATION CLOCKED .500 SECONDS
COMPONENT X-POS Z-PCS WEIGHT [oz] I I. 1250 2.6250 7.5000

2 4.0000 2.6350 3.2300

3 6.0000 0.6250 0.9500

4 6.0000 2.2800 2.0000

5 18.9400 2.4400 1.2000

6 9.0000 3.0000 2.0000

7 14.8340 2.2500 5.2200

8 12.&200 0.7500 10.3600

9 33.6700 4.2500 0.4574

10 35.5800 4.0000 0.7226

11 9.0000 -3.0000 2.4000

12 9.5000 1.5000 5.4600

X-POSITION OF THE C.g. = 9.6998

Z-POSITION OF THE C.O. = I.b286

TOTAL WEIGHT = 41.5000

MOMENT A30UT Y-AXIS {ly} = 0.0251

INCHES

INCHES

OZ

SLUg*FT**2
NOMENCLATURE

AETC

AR

O_

b

c

CD

CDo

CG

CL

Cl

CLat

CLay

CLo

CLw

Cm

Cmacw

Cmcx

Cm0ff

Cmc_t

Cm_w

Cmo

Cmof

Cmot

Cmow

Cmse

c.

Cnsr

Cp

Aerospace Engineers for the Twenty-First Century

Aspect Ratio

Angle of Attack

Wing Span

Propeller Blade Angle

Yaw Angle

Wing Mean Chord

Drag Coefficient

Parasite Drag Coefficient

Center of Gravity

Lift Coefficient

Roll Moment Coefficient

Change in Lift Coefficient with Angle of Attack for the Tail

Change in Lift Coefficient with Angle of Attack for the Vertical Tail

Lift Coefficient at Zero Angle of Attack

Change in Lift Coefficient with Angle of Attack

Roll Moment due to Yaw

Wing Lift Coefficient

Pitch Moment Coefficient

Moment Coefficient About the Aerodynamic Center of the Wing

Change in Moment Coefficient with Angle of Attack

Change in Moment Coefficient for the Fuselage

Change in Moment Coefficient for the Tail

Change in Moment Coefficient for the Wing

Total Zero Lift Moment Coefficient

Zero Lift Moment Coefficient for the Fuselage

Zero Lift Moment Coefficient for the Tail

Zero Lift Moment Coefficient for the Wing

Change in Moment Coefficient with Rudder Deflection

Yaw Moment Coefficient

Yaw Moment Coefficient due to Yaw Angle b

Rudder Control Power

Propeller Power Coefficient
Ct

e

_o

F

it

iw

J

J4D

lh

lv

mahr

n

PA

PR

Re

SFSM

SH

Sv

SW

"C

VH

VV

W

Xac

XNP

Propeller Thrust Coefficient

efficiency factor

Downwash Angle at Zero Angle of Attack

Dihedral

Tail Incidence Angle

Wing Incidence Angle

Advance Ratio

SCREEM-J4D aircraft

Horizontal Tail Moment Arm

Vertical Tail Moment Arm

Milliamp-hours

Load Factor

Efficiency Factor

Power Available

Power Required

Reynolds Number

Stick Fixed Static Margin

Horizontal Tail Surface Area

Vertical Tail Surface Area

Wing Surface Area

Flap Effectiveness

Volume Ratio of Horizontal Tail

Volume Ratio of Vertical Tail

Weight

Distance From Leading Edge to Aerodynamic Center of Wing

Distance From Wing Aerodynamic Center to Center of Gravity

Stick Fixed Neutral Point